High power epicyclic gearbox and operation thereof

ABSTRACT

A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft. A fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35 m.

This is a Continuation of application Ser. No. 17/066,713 filed Oct. 9,2020, which in turn is a Continuation of application Ser. No. 16/821,227filed Mar. 17, 2020, which in turn claims the benefit of GB 1917773.2filed Dec. 5, 2019. The disclosure of the prior applications is herebyincorporated by reference herein in its entirety.

The present disclosure relates to gas turbine engines, specifically gasturbine engines for aircraft. Aspects of the present disclosure relateto an aircraft comprising the gas turbine engine, and a method ofoperating the gas turbine engine.

Gas turbine engines for aircraft propulsion have many design factorsthat affect the overall efficiency and power output or thrust. A generalaim for a gas turbine engine is to provide low specific fuel consumption(SFC). To enable a higher thrust at a high efficiency, a larger diameterfan may be used. In order to facilitate use of a larger fan size, agearbox is provided having an output to a fan shaft via which the fan isdriven. The gearbox receives drive from a core shaft connected to aturbine system of the engine core. The gearbox allows the fan to operateat a reduced rotational speed compared to if a direct drive were used.

When making an engine having a larger fan diameter however, simplyscaling up components of a known engine type may not lead to anefficient design. For example, there may be problems associated withmounting the fan shaft within the engine. Consideration of theproperties of components used to mount the fan shaft, the properties ofthe fan shaft itself, and the properties of the gearbox components aretherefore required.

According to a first aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, the fan having a fan axial centreline; a gearbox thatreceives an input from the core shaft and outputs drive to a fan shaftvia an output of the gearbox so as to drive the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the engine, the fan shaftmounting structure comprising at least two supporting bearings connectedto the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system radial bending stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting radial bending stiffness ratio of:

$\frac{{the}{system}{radial}{bending}{stiffness}}{{the}{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.0×10⁻³.

The fan shaft mounting radial bending stiffness ratio may be greaterthan or equal to 5.0×10⁻³. The fan shaft mounting radial bendingstiffness ratio may be greater than or equal to 2.0×10⁻². The fan shaftmounting radial bending stiffness ratio may be in the range from1.0×10⁻³ to 4.0×10⁻¹. The fan shaft mounting radial bending stiffnessratio may be in the range from 5.0×10⁻³ to 1.5×10⁻¹. The fan shaftmounting radial bending stiffness ratio may be in the range from5.0×10⁻³ to 2.0×10⁻². The fan shaft mounting radial bending stiffnessratio may be in the range from 2.0×10⁻² to 1.5×10⁻¹.

The system radial bending stiffness may be greater than or equal to3.90×10⁶ N/m. The system radial bending stiffness may be greater than orequal to 3.6×10⁷ N/m. The system radial bending stiffness may be in therange from 3.90×10⁶ N/m to 1.40×10⁹ N/m. The system radial bendingstiffness may be in the range from 3.6×10⁷ N/m to 6.8×10⁸ N/m.

The radial bending stiffness of the fan shaft mounting structure may begreater than or equal to 7.00×10⁸ N/m. The radial bending stiffness ofthe fan shaft mounting structure may be greater than or equal to1.25×10⁹ N/m. The radial bending stiffness of the fan shaft mountingstructure may be in the range from 7.00×10⁸ N/m to 6.00×10¹¹ N/m. Theradial bending stiffness of the fan shaft mounting structure may be inthe range from 1.25×10⁹ N/m to 2.0×10¹¹ N/m.

The radial bending stiffness of the fan shaft at the output of thegearbox may be greater than or equal to 4.00×10⁶ N/m. The radial bendingstiffness of the fan shaft at the output of the gearbox may be greaterthan or equal to 3.7×10⁷ N/m. The radial bending stiffness of the fanshaft at the output of the gearbox may be in the range from 4.00×10⁶ N/mto 1.5×10⁹ N/m. The radial bending stiffness of the fan shaft at theoutput of the gearbox may be in the range from 3.7×10⁷ N/m to 1.0×10⁹N/m.

The product of the system radial bending stiffness and the radialbending stiffness of the fan shaft mounting structure may be greaterthan or equal to 2.7×10⁵ (N/m)². The product of the system radialbending stiffness and the radial bending stiffness of the fan shaftmounting structure may be greater than or equal to 4.0×10¹⁵ (N/m)². Theproduct of the system radial bending stiffness and the radial bendingstiffness of the fan shaft mounting structure may be in the range from2.7×10¹⁵ (N/m)² to 9.0×10¹⁹ (N/m)². The product of the system radialbending stiffness and the radial bending stiffness of the fan shaftmounting structure may be in the range from 4.0×10¹⁵ (N/m)² to 1.5×10¹⁹(N/m)².

A system tilt stiffness may be defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and fan shaft mounting tilt stiffness ratio of:

$\frac{{the}{system}{tilt}{stiffness}}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   may be greater than or equal to 1.5×10⁻³. The fan shaft mounting        tilt stiffness ratio may be greater than or equal to 6.0×10⁻³.        The fan shaft mounting tilt stiffness ratio may be greater than        or equal to 2.5×10⁻². The fan shaft mounting tilt stiffness        ratio may be in the range from 1.5×10⁻³ to 5.0×10⁻¹. The fan        shaft mounting tilt stiffness ratio may be in the range from        6.0×10⁻³ to 2.0×10⁻¹. The fan shaft mounting tilt stiffness        ratio may be in the range from 6.0×10⁻³ to 2.5×10⁻². The fan        shaft mounting tilt stiffness ratio may be in the range from        2.5×10⁻² to 2.0×10⁻.

According to a second aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, the fan having a fan axial centreline; a gearbox thatreceives an input from the core shaft and outputs drive to a fan shaftvia an output of the gearbox so as to drive the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the engine, the fan shaftmounting structure comprising at least two supporting bearings connectedto the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system tilt stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting tilt stiffness ratio of:

$\frac{{the}{system}{tilt}{stiffness}}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.5×10⁻³.

The fan shaft mounting tilt stiffness ratio may be greater than or equalto 6.0×10⁻³. The fan shaft mounting tilt stiffness ratio may be greaterthan or equal to 2.5×10⁻². The fan shaft mounting tilt stiffness ratiomay be in the range from 1.5×10⁻³ to 5.0×10⁻¹. The fan shaft mountingtilt stiffness ratio may be in the range from 6.0×10⁻³ to 2.0×10⁻¹. Thefan shaft mounting tilt stiffness ratio may be in the range from6.0×10⁻³ to 2.5×10⁻². The fan shaft mounting tilt stiffness ratio may bein the range from 2.5×10⁻² to 2.0×10⁻¹.

The system tilt stiffness may be greater than or equal to 1.10×10⁵Nm/rad. The system tilt stiffness may be greater than or equal to8.5×10⁵ Nm/rad. The system tilt stiffness may be in the range from1.10×10⁵ Nm/rad to 6.80×10⁷ Nm/rad. The system tilt stiffness may be inthe range from 8.5×10⁵ Nm/rad to 3.4×10⁷ Nm/rad.

The tilt stiffness of the fan shaft mounting structure may be greaterthan or equal to 1.50×10⁷ Nm/rad. The tilt stiffness of the fan shaftmounting structure may be greater than or equal to 2.1×10⁷ Nm/rad. Thetilt stiffness of the fan shaft mounting structure may be in the rangefrom 1.5×10⁷ Nm/rad to 2.70×10¹⁰ Nm/rad. The tilt stiffness of the fanshaft mounting structure may be in the range from 2.1×10⁷ Nm/rad to1×10¹⁰ Nm/rad.

The tilt stiffness of the fan shaft at the output of the gearbox may begreater than or equal to 7.00×10⁴ Nm/rad. The tilt stiffness of the fanshaft at the output of the gearbox may be greater than or equal to9.5×10⁵ Nm/rad. The tilt stiffness of the fan shaft at the output of thegearbox may be in the range from 7.00×10⁴ Nm/rad to 7.00×10⁷ Nm/rad. Thetilt stiffness of the fan shaft at the output of the gearbox may be inthe range from 9.5×10⁵ Nm/rad to 3.5×10⁷ Nm/rad.

The product of the system tilt stiffness and the tilt stiffness of thefan shaft mounting structure may be greater than or equal to 1.7×10¹²(Nm/rad)². The product of the system tilt stiffness and the tiltstiffness of the fan shaft mounting structure may be greater than orequal to 1.6×10¹³ (Nm/rad)². The product of the system tilt stiffnessand the tilt stiffness of the fan shaft mounting structure may be in therange from 1.7×10¹² (Nm/rad)² to 3.0×10¹⁷ (Nm/rad)². The product of thesystem tilt stiffness and the tilt stiffness of the fan shaft mountingstructure may be in the range from 1.6×10¹³ (Nm/rad)² to 3.0×10¹⁶(Nm/rad)².

One or more of the following features may apply to either or both of thefirst and second aspects above:

The fan shaft may be defined as the torque transfer component extendingfrom the output of the gearbox to the input to the fan. The fan shaftmay comprise at least part of a gearbox output shaft and at least partof a fan input shaft.

The input to the fan may be a fan input position defined as a point onthe fan shaft at the axial midpoint of the interface between the fan andthe fan shaft.

The output of the gearbox may be defined as the point of connectionbetween the fan shaft and the gearbox. The gearbox may be in a starconfiguration and the output of the gearbox may be a gearbox outputposition defined as the point of connection between the ring gear andthe fan shaft. Alternatively, the gearbox may be in a planetaryconfiguration and the output of the gearbox may be a gearbox outputposition at the interface between the fan shaft and the planet carrier.

The at least two supporting bearings may comprise a first supportingbearing and second supporting bearing.

Both of the first and the second supporting bearings may be located atpositions forward of the gearbox. Alternatively, the first supportingbearing may be located at a position forward of the gearbox and thesecond supporting bearing may be located at a position rearward of thegearbox.

The fan shaft mounting structure may further comprise a third supportingbearing. The third supporting bearing may be located between the fan andthe gearbox. The fan shaft may comprise a gearbox output shaft forming arelatively flexible portion of the fan shaft, and the fan shaft mountingstructure may comprises a gearbox output shaft support structure havingat least one gearbox output shaft bearing arranged to support thegearbox output shaft. The fan shaft mounting structure may furthercomprise one or more non-supporting softly mounted bearings. Any one ormore of the bearings provided as part of the fan shaft mountingstructure may be double bearings.

The axial distance, d₁, between the input to the fan and the closestbearing of the at least two supporting bearings in a rearward directionfrom the fan may be greater than or equal to 0.12 m. The axial distanced₁ may be greater than or equal to 0.13 m. The axial distance d₁ may bein the range from 0.12 m to 0.40 m. The axial distance d₁ may be in therange from 0.13 m to 0.30 m.

The axial distance, d₂, between the output of the gearbox and theclosest bearing of the at least two supporting bearings in a forwarddirection from the gearbox may be greater than or equal to 0.15 m. Theaxial distance d₂ may be greater than or equal to 0.16 m. The axialdistance d₂ may be in the range from 0.15 m to 0.45 m. The axialdistance d₂ may be in the range from 0.16 m to 0.40 m.

The fan-gearbox axial distance may be greater than or equal to 0.37 m.The fan-gearbox axial distance may be in the range from 0.35 m to 0.8 m.The fan-gearbox axial distance may be in the range from 0.37 m to 0.75m.

The gearbox may be an epicyclic gearbox comprising a sun gear, aplurality of planet gears, a ring gear, and a planet carrier arranged tohave the plurality of planet gears mounted thereon.

According to a third aspect there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades, thefan having a fan axial centreline; a gearbox; a power unit for drivingthe fan via the gearbox, wherein the gearbox is arranged to receive aninput from the power unit via a core shaft and output drive to a fanshaft via an output of the gearbox so as to drive the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the propulsor, the fan shaftmounting structure comprising at least two supporting bearings connectedto the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system radial bending stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting radial bending stiffness ratio of:

$\frac{{the}{system}{radial}{bending}{stiffness}}{{the}{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.0×10⁻³.

The propulsor of the third aspect may have some or all of the featuresdescribed above with respect to the gas turbine engine of the firstaspect, and may be a gas turbine engine in some embodiments.

According to a fourth aspect there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades, thefan having a fan axial centreline; a gearbox; a power unit for drivingthe fan via the gearbox, wherein the gearbox is arranged to receive aninput from the power unit via the core shaft and output drive to a fanshaft via an output of the gearbox so as to drive the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the propulsor, the fan shaftmounting structure comprising at least two supporting bearings connectedto the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system tilt stiffness is defined as

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting tilt stiffness ratio of:

$\frac{{the}{system}{tilt}{stiffness}}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.5×10⁻³.

The propulsor of the fourth aspect may have some or all of the featuresdescribed above with respect to the gas turbine engine of the secondaspect, and may be a gas turbine engine in some embodiments.

The third and fourth aspects may be combined. In such an aspect, thereis provided a propulsor for an aircraft, comprising: a fan comprising aplurality of fan blades, the fan having a fan axial centreline; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft via an output of the gearbox so asto drive the fan at a lower rotational speed than the core shaft; and afan shaft mounting structure arranged to mount the fan shaft within thepropulsor, the fan shaft mounting structure comprising at least twosupporting bearings connected to the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system radial bending stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{radial}{bending}} \\{{stiffness}{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}} \\{{the}{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting radial bending stiffness ratio of:

$\frac{{the}{system}{radial}{bending}{stiffness}}{{the}{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.0×10⁻³; and/or    -   a system tilt stiffness is defined as

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting tilt stiffness ratio of:

$\frac{{the}{system}{tilt}{stiffness}}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.5×10⁻³. The propulsor of this        aspect may have some or all of the features described above with        respect to the gas turbine engine of the first and second        aspect, and may be a gas turbine engine in some embodiments.

According to a fifth aspect there is provided a method of operating agas turbine engine for an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades, the fanhaving a fan axial centreline; a gearbox that receives an input from thecore shaft and outputs drive to a fan shaft via an output of the gearboxso as to drive the fan at a lower rotational speed than the core shaft;and a fan shaft mounting structure arranged to mount the fan shaftwithin the engine, the fan shaft mounting structure comprising at leasttwo supporting bearings connected to the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system radial bending stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{radial}{bending}} \\{{stiffness}{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}} \\{{the}{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting radial bending stiffness ratio of:

$\frac{{the}{system}{radial}{bending}{stiffness}}{{the}{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

is greater than or equal to 3.9×10⁶. The method comprises operating thegas turbine engine to provide propulsion for the aircraft under cruiseconditions.

The method of the fifth aspect may be a method of operating the gasturbine engine or the propulsor of the first aspect or third aspectrespectively. Any of the features, ratios and parameters introducedabove in connection with the first aspect or third aspect may alsotherefore apply to the fifth aspect.

According to a sixth aspect there is provided a method of operating agas turbine engine for an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades, the fanhaving a fan axial centreline; a gearbox that receives an input from thecore shaft and outputs drive to a fan shaft via an output of the gearboxso as to drive the fan at a lower rotational speed than the core shaft;and a fan shaft mounting structure arranged to mount the fan shaftwithin the engine, the fan shaft mounting structure comprising at leasttwo supporting bearings connected to the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system tilt stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting tilt stiffness ratio of:

$\frac{{the}{system}{tilt}{stiffness}}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   is greater than or equal to 1.5×10⁻³. The method comprises        operating the gas turbine engine to provide propulsion for the        aircraft under cruise conditions.

The method of the sixth aspect may be a method of operating the gasturbine engine or the propulsor of the second aspect or fourth aspectsrespectively. Any of the features, ratios and parameters introducedabove in connection with the second aspect or fourth aspect may alsotherefore apply to the sixth aspect.

The fifth and sixth aspects may be combined. Such an aspect may providea method of operating a gas turbine engine for an aircraft, the gasturbine engine comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, the fan having a fan axial centreline; a gearbox thatreceives an input from the core shaft and outputs drive to a fan shaftvia an output of the gearbox so as to drive the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the engine, the fan shaftmounting structure comprising at least two supporting bearings connectedto the fan shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system radial bending stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{radial}{bending}} \\{{stiffness}{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}} \\{{the}{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting radial bending stiffness ratio of:

$\frac{{the}{system}{radial}{bending}{stiffness}}{{the}{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

is greater than or equal to 3.9×10⁶; and/or

-   -   a system tilt stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

-   -   and a fan shaft mounting tilt stiffness ratio of:

$\frac{{the}{system}{tilt}{stiffness}}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

is greater than or equal to 1.5×10⁻³. The method comprises operating thegas turbine engine to provide propulsion for the aircraft under cruiseconditions. This aspect may be a method of operating the gas turbine orpropulsor of any preceding aspect.

The inventor has discovered that designing the fan shaft and fan shaftmounting structure so that the fan shaft mounting radial bending/tiltstiffness ratio is within the specified range allows for improvedisolation from loads transmitted from the fan while maintainingsufficient location of the fan within the engine. The inventor has foundthat if the stiffness of the fan shaft mounting structure is reduced sothat the fan shaft mounting stiffness ratio is outside of the specifiedrange the fan will not be adequately located within the engine, leadingto problems with fan tip clearance control. For example, movement of thefan relative to its surrounding structure may lead to a higher fan tipclearance being required, which would reduce the overall efficiency ofthe engine. The inventor has also found that further increasing thestiffness of the fan shaft mounting structure so that the fan shaftmounting stiffness ratio is outside of the specified ranged there wouldbe little or no practical benefit in terms of locating the fan, butwould instead lead to undesirable increases in weight of the mountingstructure. The inventor has found that increasing the stiffness of thefan shaft so that the fan shaft mounting stiffness ratio is outside ofthe specified limit would result in excessive loads being transmitted tothe gearbox from the fan. While decreasing the fan shaft stiffness istherefore advantageous, the inventor has found that further reducing thestiffness of the fan shaft so that the fan shaft mounting stiffnessratio is outside of the range above would cause undesired low modallateral vibrations in the fan shaft with excessive amplitude.

In other aspects, value ranges for the product of the components of thefan shaft mounting radial bending/tilt stiffness ratios may be specifiedinstead of, or as well as, value ranges for the ratios.

According to one such aspect, the first aspect introduced above may beformulated as an aspect providing a gas turbine engine for an aircraftcomprising: an engine core comprising a turbine, a compressor, and acore shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades, the fan having a fan axial centreline; a gearbox that receivesan input from the core shaft and outputs drive to a fan shaft via anoutput of the gearbox so as to drive the fan at a lower rotational speedthan the core shaft; and a fan shaft mounting structure arranged tomount the fan shaft within the engine, the fan shaft mounting structurecomprising at least two supporting bearings connected to the fan shaft,and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system radial bending stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{radial}{bending}} \\{{stiffness}{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}} \\{{the}{output}{of}{the}{gearbox}}\end{matrix}} )};$

and the product of the system radial bending stiffness and the radialbending stiffness of the fan shaft mounting structure may be may begreater than or equal to 2.7×10¹⁵ (N/m)², greater than or equal to4.0×10¹⁵ (N/m)², in the range from 2.7×10¹⁵ (N/m)² to 9.0×10¹⁹ (N/m)²,or in the range from 4.0×10¹⁵ (N/m)² to 1.5×10¹⁹ (N/m)².

According to another such aspect, the second aspect introduced above maybe formulated as an aspect providing a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades, the fan having a fan axial centreline; a gearbox that receivesan input from the core shaft and outputs drive to a fan shaft via anoutput of the gearbox so as to drive the fan at a lower rotational speedthan the core shaft; and a fan shaft mounting structure arranged tomount the fan shaft within the engine, the fan shaft mounting structurecomprising at least two supporting bearings connected to the fan shaft,and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the fan axial centreline,        the fan-gearbox axial distance being greater than or equal to        0.35 m;    -   a system tilt stiffness is defined as:

$\frac{1}{( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}} \\{{mounting}{structure}}\end{matrix}} ) + ( \frac{1}{\begin{matrix}{{the}{tilt}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}} \\{{output}{of}{the}{gearbox}}\end{matrix}} )};$

and the product of the system tilt stiffness and the tilt stiffness ofthe fan shaft mounting structure may be greater than or equal to1.7×10¹² (Nm/rad)², greater than or equal to 1.6×10¹³ (Nm/rad)², in therange from 1.7×10¹² (Nm/rad)² to 3.0×10¹⁷ (Nm/rad)², or in the rangefrom 1.6×10¹³ (Nm/rad)² to 3.0×10¹⁶ (Nm/rad)²

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly.

According to a seventh aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades; a gearbox that receives an input from the core shaft andoutputs drive to a fan shaft via an output of the gearbox so as to drivethe fan via an input to the fan at a lower rotational speed than thecore shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m, and    -   a fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{system}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 6.0×10⁻³.

The fan shaft radial bending stiffness ratio may be greater than orequal to 0.015. The fan shaft radial bending stiffness ratio may be inthe range from 6.0×10⁻³ to 2.5×10¹. The fan shaft radial bendingstiffness ratio may be in the range from 0.015 to 2.5.

The radial bending stiffness of the fan shaft at the input to the fanmay be greater than or equal to 3.00×10⁶ N/m. The radial bendingstiffness of the fan shaft at the input to the fan may be greater thanor equal to 6.3×10⁶ N/m. The radial bending stiffness of the fan shaftat the input to the fan may be in the range from 3.00×10⁶ N/m to2.00×10⁹ N/m. The radial bending stiffness of the fan shaft at the inputto the fan may be in the range from 6.3×10⁶ N/m to 1.0×10⁹ N/m.

The radial bending stiffness of the fan shaft at the output of thegearbox may be greater than or equal to 4.00×10⁶ N/m. The radial bendingstiffness of the fan shaft at the output of the gearbox may be greaterthan or equal to 3.7×10⁷ N/m. The radial bending stiffness of the fanshaft at the output of the gearbox may be in the range from 4.00×10⁶ N/mto 1.5×10⁹ N/m. The radial bending stiffness of the fan shaft at theoutput of the gearbox may be in the range from 3.7×10⁷ N/m to 1.0×10⁹N/m.

The diameter of the fan may be in the range from 240 cm to 280 cm. Insuch an embodiment, the fan shaft radial bending stiffness ratio may begreater than or equal to 0.03, or in the range from 0.03 to 0.85.

The diameter of the fan may be in the range from 330 cm to 380 cm. Insuch an embodiment, the fan shaft radial bending stiffness ratio may begreater than or equal to 0.02 or in the range from 0.02 to 1.5.

A fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

may be greater than or equal to 2.5×10⁻².

The fan shaft tilt stiffness ratio may be greater than or equal to 0.05.The fan shaft tilt stiffness ratio may be in the range from 2.5×10⁻² to3.7×10². The fan shaft tilt stiffness ratio may be in the range from0.05 to 4.0×10¹.

A product of:

(the radial bending stiffness of the fan shaft at the input to thefan)×(the radial bending stiffness of the fan shaft at the output of thegearbox)

may be greater than or equal to 1.2×10¹³ (N/m)², greater than or equalto 2.4×10¹⁴ (N/m)², in the range from 1.2×10¹³ (N/m)² to 3.0×10¹⁸(N/m)², or in the range from 2.4×10¹⁴ (N/m)² to 3.0×10¹⁷ (N/m)².

According to an eighth aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades; a gearbox that receives an input from the core shaft andoutputs drive to a fan shaft via an output of the gearbox so as to drivethe fan via an input to the fan at a lower rotational speed than thecore shaft, wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35, and    -   a fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 2.5×10⁻².

The fan shaft tilt stiffness ratio may be greater than or equal to 0.05.The fan shaft tilt stiffness ratio may be in the range from 2.5×10⁻² to3.7×10². The fan shaft tilt stiffness ratio may be in the range from0.05 to 4.0×10¹.

The tilt stiffness of the fan shaft at the input to the fan may begreater than or equal to 5.00×10⁵ Nm/rad. The tilt stiffness of the fanshaft at the input to the fan may be greater than or equal to 9.0×10⁵Nm/rad. The tilt stiffness of the fan shaft at the input to the fan maybe in the range from 5.00×10⁵ Nm/rad to 7.00×10⁸ Nm/rad. The tiltstiffness of the fan shaft at the input to the fan may be in the rangefrom 9.0×10⁵ Nm/rad to 3.5×10⁸ Nm/rad.

The tilt stiffness of the fan shaft at the output of the gearbox may begreater than or equal to 7.00×10⁴ Nm/rad. The tilt stiffness of the fanshaft at the output of the gearbox may be greater than or equal to9.5×10⁵ Nm/rad. The tilt stiffness of the fan shaft at the output of thegearbox may be in the range from 7.00×10⁴ Nm/rad to 7.00×10⁷ Nm/rad. Thetilt stiffness of the fan shaft at the output of the gearbox may be inthe range from 9.5×10⁵ Nm/rad to 3.5×10⁷ Nm/rad.

The diameter of the fan may be in the range from 240 cm to 280 cm. Insuch an embodiment, the fan shaft tilt stiffness ratio may be greaterthan or equal to 0.2 or in the range from 0.2 to 5.0.

The diameter of the fan may be in the range from 330 cm to 380 cm. Insuch an embodiment, the fan shaft tilt stiffness ratio may be greaterthan or equal to 0.1 or in the range from 0.1 to 1.0×10¹.

A fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

may be greater than or equal to 6.0×10⁻³. The fan shaft radial bendingstiffness ratio may be greater than or equal to 0.015. The fan shaftradial bending stiffness ratio may be in the range from 6.0×10⁻³ to2.5×10¹. The fan shaft radial bending stiffness ratio may be in therange from 0.015 to 2.5.

A product of:

(the tilt stiffness of the fan shaft at the input to the fan)×(the tiltstiffness of the fan shaft at the output of the gearbox)

may be greater than or equal to 3.5×10¹⁰ (Nm/rad)², greater than orequal to 7.2×10¹¹ (Nm/rad)², in the range from 3.5×10¹⁰ (Nm/rad)² to5.0×10¹⁶ (Nm/rad)², or in the range from 7.2×10¹¹ (Nm/rad)² to 5.0×10¹⁵(Nm/rad)² One or more of the following features may apply to either orboth of the preceding two aspects (e.g. the seventh and eighth aspects):

The fan-gearbox axial distance may be greater than or equal to 0.37 m.The fan-gearbox axial distance may be in the range from 0.35 m to 0.8 m.The fan-gearbox axial distance may be in the range from 0.37 m to 0.75m.

The fan shaft may be defined as the torque transfer component extendingfrom the output of the gearbox to the input to the fan. The fan shaftmay comprise at least part (or all) of a gearbox output shaft and atleast part (or all) of a fan input shaft.

The input to the fan may be a fan input position defined as a point onthe fan shaft at the axial midpoint of the interface between the fan andthe fan shaft.

The output of the gearbox may be defined as the point of connectionbetween the fan shaft and the gearbox. The gearbox may be in a starconfiguration and the output of the gearbox may be a gearbox outputposition defined as the point of connection between the ring gear andthe fan shaft. The gearbox may be in a planetary configuration and theoutput of the gearbox may be a gearbox output position at the interfacebetween the fan shaft and the planet carrier.

The gas turbine engine may further comprise a fan shaft mountingstructure arranged to mount the fan shaft within the engine. The fanshaft mounting structure may comprise at least two supporting bearingsconnected to the fan shaft.

The at least two supporting bearings may comprise a first supportingbearing and second supporting bearing. Both of the first and the secondsupporting bearings may be located at positions forward of the gearbox.Alternatively, the first supporting bearing may be located at a positionforward of the gearbox and the second supporting bearing may be locatedat a position rearward of the gearbox.

The fan shaft mounting structure may comprise a third supportingbearing. The third supporting bearing may be located between the fan andthe gearbox.

The fan shaft may comprise a gearbox output shaft forming a relativelyflexible portion of the fan shaft. The fan shaft mounting structure maycomprise gearbox output shaft support structure having at least onegearbox output shaft bearing arranged to support the gearbox outputshaft.

The fan shaft mounting structure may further comprise one or morenon-supporting softly mounted bearings.

Any one or more of the bearings provided as part of the fan shaftmounting structure may be double bearings.

The axial distance, d₁, between the input to the fan and the closestbearing of the at least two supporting bearings in a rearward directionfrom the fan may be greater than or equal to 0.12 m, greater than orequal to 0.13 m, in the range from 0.12 m to 0.40 m, or in the rangefrom 0.13 m to 0.30 m.

The axial distance, d₂, between the output of the gearbox and theclosest bearing of the at least two supporting bearings in a forwarddirection from the gearbox may be greater than or equal to 0.15 m,greater than or equal to 0.16 m, in the range from 0.15 m to 0.45 m, orin the range from 0.16 m to 0.40 m.

The gearbox may be an epicyclic gearbox comprising a sun gear, aplurality of planet gears, a ring gear, and a planet carrier arranged tohave the plurality of planet gears mounted thereon.

According to a ninth aspect there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft via an output of the gearbox so asto drive the fan via an input to the fan at a lower rotational speedthan the core shaft, and wherein

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m; and wherein:    -   a fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 6.0×10⁻³.

The propulsor of the ninth aspect may have some or all of the featuresdescribed above with respect to the gas turbine engine of the seventhaspect, and may be a gas turbine engine in some embodiments.

According to a tenth aspect there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft via an output of the gearbox so asto drive the fan via an input to the fan at a lower rotational speedthan the core shaft, and wherein

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m; and wherein:    -   a fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 2.5×10⁻².

The propulsor of the tenth aspect may have some or all of the featuresdescribed above with respect to the gas turbine engine of the eighthaspect, and may be a gas turbine engine in some embodiments.

The ninth and tenth aspect may be combined. According to an eleventhaspect, there is provided a propulsor for an aircraft, comprising: a fancomprising a plurality of fan blades; a gearbox; a power unit fordriving the fan via the gearbox, wherein the gearbox is arranged toreceive an input from the power unit via a core shaft and output driveto a fan shaft via an output of the gearbox so as to drive the fan viaan input to the fan at a lower rotational speed than the core shaft, andwherein

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m; and wherein:    -   a) a fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 6.0×10⁻³; and/or    -   b) a fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 2.5×10⁻².

The propulsor of the eleventh aspect may have some or all of thefeatures described above with respect to the gas turbine engine of theseventh or eighth aspect, and may be a gas turbine engine in someembodiments.

According to a twelfth aspect there is provided a method of operating agas turbine engine for an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades; a gearboxthat receives an input from the core shaft and outputs drive to a fanshaft via an output of the gearbox so as to drive the fan via an inputto the fan at a lower rotational speed than the core shaft, and wherein

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m, wherein:    -   a fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

is greater than or equal to 6.0×10⁻³. The method comprises operating thegas turbine engine to provide propulsion for the aircraft under cruiseconditions.

The method of the twelfth aspect may be a method of operating the gasturbine engine or the propulsor of the seventh aspect or ninth aspectrespectively. Any of the features, ratios and parameters introducedabove in connection with the seventh or ninth aspect may also thereforeapply to the twelfth aspect.

According to a thirteenth aspect there is provided a method of operatinga gas turbine engine for an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, and a core shaftconnecting the turbine to the compressor; a fan located upstream of theengine core, the fan comprising a plurality of fan blades; a gearboxthat receives an input from the core shaft and outputs drive to a fanshaft via an output of the gearbox so as to drive the fan via an inputto the fan at a lower rotational speed than the core shaft, and wherein

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m, wherein:    -   a fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 2.5×10⁻². The method comprises        operating the gas turbine engine to provide propulsion for the        aircraft under cruise conditions.

The method of the thirteenth aspect may be a method of operating the gasturbine engine or the propulsor of the eighth aspect or tenth aspectrespectively. Any of the features, ratios and parameters introducedabove in connection with the eighth aspect or tenth aspect may alsotherefore apply to the thirteenth aspect.

The twelfth and thirteenth aspects may be combined. According to afourteenth aspect, there is provided a method of operating a gas turbineengine for an aircraft, the gas turbine engine comprising: an enginecore comprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; a gearbox that receives aninput from the core shaft and outputs drive to a fan shaft via an outputof the gearbox so as to drive the fan via an input to the fan at a lowerrotational speed than the core shaft, and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m, wherein:    -   a) a fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 6.0×10⁻³; and/or    -   b) a fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

-   -   is greater than or equal to 2.5×10⁻², the method comprising        operating the gas turbine engine to provide propulsion for the        aircraft under cruise conditions.

The method of the fourteenth aspect may be a method of operating the gasturbine engine or the propulsor of the seventh, eighth or eleventhaspect. Any of the features, ratios and parameters introduced above inconnection with the seventh, eighth or eleventh aspect may alsotherefore apply to the fourteenth aspect.

The inventor has discovered that by designing the comparative stiffnessof the fan shaft at the fan input compared to the gearbox output, withinother constraints of the engine, the fan can be accurately locatedwithin the engine while also isolating the gearbox from loadstransmitted from the fan. The inventor has found that if the stiffnessof the fan shaft at the fan input is such that the fan shaft radialbending/tilt stiffness ratio is below the specified range the fan wouldnot be adequately located (leading to fan tip control problems) andthere would be undesirable low modal vibrations with high amplitude. Theinventor has found that if the fan shaft is designed so that thestiffness at the fan input was further increased so that the fan shaftradial bending/tilt stiffness ratio is outside of the above range therewould be little or no practical benefit in terms of improved location ofthe fan without undesirable increases in overall weight. The inventorhas also found that increasing the stiffness of the fan shaft at theoutput of the gearbox so that the fan shaft radial bending/tiltstiffness ratio is outside of the specified range would lead to excessload transmission into the gearbox from the fan. The inventor has alsofound that reducing the stiffness of the fan shaft at the gearbox outputso that the fan shaft radial bending/tilt stiffness ratio is outside ofthe range above would cause undesirable low modal lateral vibrations inthe fan shaft with excessive amplitude.

In other aspects, value ranges for the product of the components of thefan shaft radial bending stiffness ratio and the fan shaft tiltstiffness ratio may be specified instead of, or as well as, value rangesfor the ratios.

According to one such aspect, the seventh aspect introduced above may beformulated as an aspect providing a gas turbine engine for an aircraftcomprising: an engine core comprising a turbine, a compressor, and acore shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; a gearbox that receives an input from the core shaft and outputsdrive to a fan shaft via an output of the gearbox so as to drive the fanvia an input to the fan at a lower rotational speed than the core shaft,and wherein:

-   -   a fan-gearbox axial distance is defined as the axial distance        between the output of the gearbox and the axial centreline of        the fan, the fan-gearbox axial distance being greater than or        equal to 0.35 m, and a product of:

(the radial bending stiffness of the fan shaft at the input to thefan)×(the radial bending stiffness of the fan shaft at the output of thegearbox)

may be greater than or equal to 1.2×10¹³ (N/m)², greater than or equalto 2.4×10¹⁴ (N/m)², in the range from 1.2×10¹³ (N/m)² to 3.0×10¹⁸(N/m)², or in the range from 2.4×10¹⁴ (N/m)² to 3.0×10¹⁷ (N/m)².

According to another such aspect, the eighth aspect introduced above maybe formulated as an aspect providing a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; a gearbox that receives an input from the core shaft and outputsdrive to a fan shaft via an output of the gearbox so as to drive the fanvia an input to the fan at a lower rotational speed than the core shaft,wherein:

a fan-gearbox axial distance is defined as the axial distance betweenthe output of the gearbox and the axial centreline of the fan, thefan-gearbox axial distance being greater than or equal to 0.35, and aproduct of:

(the tilt stiffness of the fan shaft at the input to the fan)×(the tiltstiffness of the fan shaft at the output of the gearbox)

may be greater than or equal to 3.5×10¹⁰ (Nm/rad)², greater than orequal to 7.2×10¹¹ (Nm/rad)², in the range from 3.5×10¹⁰ (Nm/rad)² to5.0×10¹⁶ (Nm/rad)², or in the range from 7.2×10¹¹ (Nm/rad)² to 5.0×10¹⁵(Nm/rad)².

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly.

According to a fifteenth aspect there is provided a gas turbine enginefor an aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades; a gearbox that receives an input from the core shaft andoutputs drive via an output of the gearbox to a fan shaft so as to drivethe fan, via an input to the fan, at a lower rotational speed than thecore shaft; and

-   -   a fan shaft mounting structure arranged to mount the fan shaft        within the engine, wherein the fan shaft mounting structure        comprises at least two supporting bearings connected to the fan        shaft, wherein:    -   the output of the gearbox is at a gearbox output position and        the input to the fan is at a fan input position;    -   a first bearing separation distance, d₁, is defined as the axial        distance between the input to the fan and the closest bearing of        the at least two supporting bearings in a rearward direction        from the fan; and    -   a first bearing separation ratio of:

$\frac{{the}{first}{bearing}{seperation}{{distance}{}( d_{1} )}}{\begin{matrix}{{the}{axial}{distance}{between}{the}{fan}{input}{position}} \\{{and}{the}{gearbox}{output}{{position}{}( d_{4} )}}\end{matrix}}$

-   -   is greater than or equal to 1.6×10⁻¹, and the axial distance        between the fan input position and the gearbox output position,        d₄, is greater than or equal to 0.43 m.

The first bearing separation ratio may be greater than or equal to1.8×10⁻¹. The first bearing separation ratio may be greater than orequal to 1.6×10⁻¹. The first bearing separation ratio may be greaterthan or equal to 2.2×10⁻¹. The first bearing separation ratio may be inthe range from 1.6×10¹ to 3.3×10¹. The first bearing separation ratiomay be in the range from 1.8×10⁻¹ to 3.0×10⁻¹. The first bearingseparation ratio may be in the range from 1.6×10⁻¹ to 2.2×10⁻¹. Thefirst bearing separation ratio may be in the range from 2.2×10⁻¹ to3.3×10⁻¹.

The first bearing separation distance, d₁, may be greater than or equalto 0.12 m, greater than or equal to 0.13 m, in the range from 0.12 m to0.40 m, or in the range from 0.13 m to 0.30 m.

A second bearing separation distance, d₂, is defined as the axialdistance between the output of the gearbox and the closest bearing ofthe at least two supporting bearings in a forward direction from thegearbox. The second bearing separation distance, d₂, maybe greater thanor equal to 0.15 m, greater than or equal to 0.16 m, in the range from0.15 m to 0.45 m, or in the range from 0.16 m to 0.40 m.

The axial distance between the fan input position and the gearbox outputposition, d₄, may be greater than or equal to 0.46 m, in the range from0.43 m to 0.95 m, or in the range from 0.46 m to 0.85 m.

A first bearing separation product defined as:

the first bearing separation distance (d₁)×the axial distance betweenthe fan input position and the gearbox output position (d₄)

may be greater than or equal to 5.2×10⁻² m². The first bearingseparation product may be greater than or equal to 5.7×10⁻² m². Thefirst bearing separation product may be in the range from 5.2×10⁻² m² to2.6×10⁻¹ m². The first bearing separation product may be in the rangefrom 5.7×10⁻² m² to 2.4×10⁻¹ m.

Both of the at least two supporting bearings may be located at positionsforward of the gearbox. A bearing axial separation, d₃, may be definedas the axial distance between the supporting bearing that is the closestbearing of the at least two supporting bearings in a rearward directionfrom the fan and the supporting bearing that is the closest bearing ofthe at least two supporting bearings in a forward direction from thegearbox. A second bearing separation ratio defined as:

$\frac{{the}{first}{bearing}{seperation}{distance}( d_{1} )}{{the}{bearing}{axial}{seperation}( d_{3} )}$

may be greater than or equal to 4.1×10⁻¹. The second bearing separationratio may be greater than or equal to 4.5×10⁻¹. The second bearingseparation ratio may be greater than or equal to 6.0×10⁻¹. The secondbearing separation ratio may be in the range from 4.1×10⁻¹ to 8.3×10⁻¹.The second bearing separation ratio may be in the range from 4.5×10⁻¹ to7.7×10⁻¹. The second bearing separation ratio may be in the range from4.1×10⁻¹ to 6.0×10⁻¹. The second bearing separation ratio may in therange from 6.0×10⁻¹ to 8.3×10⁻¹.

One of the at least two supporting bearings may be located at a positionforward of the gearbox and another of the at least two supportingbearings may be located at a position rearward of the gearbox.

The at least two supporting bearings may include a first supportingbearing and a second supporting bearing, and the fan shaft mountingstructure may comprises a third supporting bearing. The third supportingbearing may be located at a position between the fan and the gearbox.

The fan shaft may comprise a gearbox output shaft forming a relativelyflexible portion of the fan shaft. The fan shaft mounting structure maycomprise a gearbox output shaft support structure having at least onegearbox output shaft bearing arranged to support the gearbox outputshaft.

The fan shaft mounting structure may further comprise one or morenon-supporting softly mounted bearings.

Any one or more of the bearings provided as part of the fan shaftmounting structure may be double bearings.

A fan shaft radial bending stiffness ratio of:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

may be greater than or equal to 6.0×10⁻³. The fan shaft radial bendingstiffness ratio may be greater than or equal to 0.015. The fan shaftradial bending stiffness ratio may be in the range from 6.0×10⁻³ to2.5×10¹. The fan shaft radial bending stiffness ratio may be in therange from 0.015 to 2.5.

The radial bending stiffness of the fan shaft at the input to the fanmay be greater than or equal to 3.00×10⁶ N/m. The radial bendingstiffness of the fan shaft at the input to the fan may be greater thanor equal to 6.3×10⁶ N/m. The radial bending stiffness of the fan shaftat the input to the fan may be in the range from 3.00×10⁶ N/m to2.00×10⁹ N/m. The radial bending stiffness of the fan shaft at the inputto the fan may be in the range from 6.3×10⁶ N/m to 1.0×10⁹ N/m.

The radial bending stiffness of the fan shaft at the output of thegearbox may be greater than or equal to 4.00×10⁶ N/m. The radial bendingstiffness of the fan shaft at the output of the gearbox may be greaterthan or equal to 3.7×10⁷ N/m. The radial bending stiffness of the fanshaft at the output of the gearbox may be in the range from 4.00×10⁶ N/mto 1.5×10⁹ N/m. The radial bending stiffness of the fan shaft at theoutput of the gearbox may be in the range from 3.7×10⁷ N/m to 1.0×10⁹N/m.

The diameter of the fan may be in the range from 240 cm to 280 cm. Insuch an embodiment, the fan shaft radial bending stiffness ratio may begreater than or equal to 0.03, or in the range from 0.03 to 0.85.

Alternatively, the diameter of the fan may be in the range from 330 cmto 380 cm. In such an embodiment, the fan shaft radial bending stiffnessratio may be greater than or equal to 0.02, or in the range from 0.02 to1.5.

A fan shaft tilt stiffness ratio of:

$\frac{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}{at}{the}{input}{to}{the}{fan}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}} \\{{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}}$

may be greater than or equal to 2.5×10⁻². The fan shaft tilt stiffnessratio may be greater than or equal to 0.05. The fan shaft tilt stiffnessratio may be in the range from 2.5×10⁻² to 3.7×10². The fan shaft tiltstiffness ratio may be in the range from 0.05 to 4.0×10¹.

The tilt stiffness of the fan shaft at the input to the fan may begreater than or equal to 5.00×10⁵ Nm/rad. The tilt stiffness of the fanshaft at the input to the fan may be greater than or equal to 9.0×10⁵Nm/rad. The tilt stiffness of the fan shaft at the input to the fan maybe in the range from 5.00×10⁵ Nm/rad to 7.00×10⁸ Nm/rad. The tiltstiffness of the fan shaft at the input to the fan may be in the rangefrom 9.0×10⁵ Nm/rad to 3.5×10⁸ Nm/rad.

The tilt stiffness of the fan shaft at the output of the gearbox may begreater than or equal to 7.00×10⁴ Nm/rad. The tilt stiffness of the fanshaft at the output of the gearbox may be greater than or equal to9.5×10⁵ Nm/rad. The tilt stiffness of the fan shaft at the output of thegearbox may be in the range from 7.00×10⁴ Nm/rad to 7.00×10⁷ Nm/rad. Thetilt stiffness of the fan shaft at the output of the gearbox may be inthe range from 9.5×10⁵ Nm/rad to 3.5×10⁷ Nm/rad.

The diameter of the fan may be in the range from 240 cm to 280 cm. Insuch an embodiment the fan shaft tilt stiffness ratio may be greaterthan or equal to 0.2 or in the range from 0.2 to 5.0.

Alternatively, the diameter of the fan may be in the range from 330 cmto 380 cm. In such an embodiment, the fan shaft tilt stiffness ratio maybe greater than or equal to 0.1 or in the range from 0.1 to 1.0×10¹.

The fan shaft may be defined as the torque transfer component extendingfrom the output of the gearbox to the input to the fan. The fan shaftmay comprise at least part of a gearbox output shaft and at least partof a fan input shaft.

The input to the fan may be a fan input position defined as a point onthe fan shaft at the axial midpoint of the interface between the fan andthe fan shaft.

The output of the gearbox may be defined as the point of connectionbetween the fan shaft and the gearbox. The gearbox may be in a starconfiguration and the output of the gearbox may be a gearbox outputposition defined as the point of connection between the ring gear andthe fan shaft. Alternatively, the gearbox may be in a planetaryconfiguration and the output of the gearbox may be a gearbox outputposition at the interface between the fan shaft and the planet carrier.

The gearbox may be an epicyclic gearbox comprising a sun gear, aplurality of planet gears, a ring gear, and a planet carrier arranged tohave the plurality of planet gears mounted thereon.

According to a sixteenth aspect, there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft so as to drive the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the propulsor, wherein the fanshaft mounting structure comprises at least two supporting bearingsconnected to the fan shaft, wherein:

-   -   the output of the gearbox is at a gearbox output position and        the input to the fan is at a fan input position;    -   a first bearing separation distance, d₁, is defined as the axial        distance between the input to the fan and the closest bearing of        the at least two supporting bearings in a rearward direction        from the fan; and    -   a first bearing separation ratio of:

$\frac{{the}{first}{bearing}{separation}{distance}( d_{1} )}{\begin{matrix}{{the}{axial}{distance}{between}{the}{fan}{input}} \\{{position}{and}{the}{gearbox}{output}{{position}{}( d_{4} )}}\end{matrix}}$

-   -   is greater than or equal to 1.6×10⁻¹, and the axial distance        between the fan input position and the gearbox output position,        d₄, is greater than or equal to 0.43 m.

The propulsor may have some or all of the features described above withrespect to the gas turbine engine of the fifteenth aspect, and may be agas turbine engine in some embodiments.

According to a seventeenth aspect, there is provided a method ofoperating a gas turbine engine for an aircraft comprising: an enginecore comprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; a gearbox that receives aninput from the core shaft and outputs drive via an output of the gearboxto a fan shaft so as to drive the fan via an input to the fan at a lowerrotational speed than the core shaft; and a fan shaft mounting structurearranged to mount the fan shaft within the engine, wherein the fan shaftmounting structure comprises at least two supporting bearings connectedto the fan shaft, wherein:

-   -   the output of the gearbox is at a gearbox output position and        the input to the fan is at a fan input position;    -   a first bearing separation distance, d₁, is defined as the axial        distance between the input to the fan and the closest bearing of        the at least two supporting bearings in a rearward direction        from the fan; and    -   a first bearing separation ratio of:

$\frac{{the}{first}{bearing}{separation}{distance}( d_{1} )}{\begin{matrix}{{the}{axial}{distance}{between}{the}{fan}{input}} \\{{position}{and}{the}{gearbox}{output}{{position}{}( d_{4} )}}\end{matrix}}$

-   -   is greater than or equal to 1.6×10⁻¹, and the axial distance        between the fan input position and the gearbox output position        (d₄) is greater than or equal to 0.43 m. The method comprises        operating the gas turbine engine to provide propulsion for the        aircraft under cruise conditions.

The method of the seventeenth aspect may be a method of operating thegas turbine engine or the propulsor of the fifteenth aspect or sixteenthaspect respectively. Any of the features, ratios and parametersintroduced above in connection with the fifteenth or sixteenth aspectalso therefore apply to the seventeenth aspect.

The inventor has discovered that by arranging the bearings so that thefirst bearing separation ratio defined above is within the specifiedrange the fan can be sufficiently located within the engine and thegearbox isolated from loads from the fan, while still providing asuitable fan shaft geometry to fit within the engine. The inventor hasfound that if the first bearing separation distance were to be increasedso that the ratio was outside the specified range the fan would not beadequately located. The inventor has also discovered that if the axialdistance between the fan input and gearbox output were to be decreasedso that the ratio was outside of the specified range there would beexcessive transmission of load into the gearbox from the fan.

In other aspects, value ranges for the product of the components of thefirst bearing separation ratio may be specified instead of, or as wellas, value ranges for the ratios.

According to one such aspect, the fifteenth aspect introduced above maybe formulated as an aspect providing a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; a gearbox that receives an input from the core shaft and outputsdrive via an output of the gearbox to a fan shaft so as to drive thefan, via an input to the fan, at a lower rotational speed than the coreshaft; and a fan shaft mounting structure arranged to mount the fanshaft within the engine, wherein the fan shaft mounting structurecomprises at least two supporting bearings connected to the fan shaft,wherein: the output of the gearbox is at a gearbox output position andthe input to the fan is at a fan input position; a first bearingseparation distance, d₁, is defined as the axial distance between theinput to the fan and the closest bearing of the at least two supportingbearings in a rearward direction from the fan; and a first bearingseparation product defined as:

the first bearing separation distance (d₁)×the axial distance betweenthe fan input position and the gearbox output position (d₄)

is greater than or equal to 5.2×10⁻² m², greater than or equal to5.7×10⁻² m², in the range from 5.2×10⁻² m² to 2.6×10⁻¹ m², or in therange from 5.7×10⁻² m² to 2.4×10⁻¹ m², and the axial distance betweenthe fan input position and the gearbox output position, d₄, is greaterthan or equal to 0.43 m.

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly.

According to a eighteenth aspect, there is provided a gas turbine enginefor an aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades; a gearbox that receives an input from the core shaft andoutputs drive to a fan shaft so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier arranged to have the plurality of planet gearsmounted thereon; and

-   -   a fan shaft mounting structure arranged to mount the fan shaft        within the engine, the fan shaft mounting structure comprising        at least two supporting bearings connected to the fan shaft, and        wherein:    -   the torsional stiffness of the planet carrier is greater than or        equal to 1.60×10⁸ Nm/rad; and    -   the radial bending stiffness of the fan shaft mounting structure        is greater than or equal to 7.00×10⁸ N/m.

The torsional stiffness of the planet carrier may be greater than orequal to 2.7×10⁸ Nm/rad. The torsional stiffness of the planet carriermay be in the range from 1.60×10⁸ Nm/rad to 1.00×10¹¹ Nm/rad. Thetorsional stiffness of the planet carrier may be in the range from2.7×10⁸ Nm/rad to 1×10¹⁰ Nm/rad.

The radial bending stiffness of the fan shaft mounting structure may begreater than or equal to 1.25×10⁹ N/m. The radial bending stiffness ofthe fan shaft mounting structure may be in the range from 7.00×10⁸ N/mto 6.00×10¹¹ N/m. The radial bending stiffness of the fan shaft mountingstructure may be in the range from 1.25×10⁹ N/m to 2.0×10¹¹ N/m.

The fan may have a fan diameter in the range from 240 cm to 280 cm. Insuch an embodiment, the torsional stiffness of the planet carrier may begreater than or equal to 1.8×10⁸ Nm/rad or in the range from 1.8×10⁸Nm/rad to 4.8×10⁹ Nm/rad. Additionally, or alternatively, in such anembodiment, the radial bending stiffness of the fan shaft mountingstructure may be greater than or equal to 7.0×10⁸ N/m, or in the rangefrom 7.0×10⁸ N/m to 5.0×10¹¹ N/m.

The fan may have a fan diameter in the range from 330 cm to 380 cm. Insuch an embodiment, the torsional stiffness of the planet carrier may begreater than or equal to 6.0×10⁸ Nm/rad or in the range from 6.0×10⁸Nm/rad to 2.2×10¹⁰ Nm/rad. Additionally, or alternatively, in such anembodiment, the radial bending stiffness of the fan shaft mountingstructure may be greater than or equal to 1.4×10⁹ N/m, or in the rangefrom 1.4×10⁹ N/m to 6.0×10¹¹ N/m.

The tilt stiffness of the fan shaft mounting structure may be greaterthan or equal to 1.50×10⁷ Nm/rad. The tilt stiffness of the fan shaftmounting structure may be greater than or equal to 2.1×10⁷ Nm/rad. Thetilt stiffness of the fan shaft mounting structure may be in the rangefrom 1.5×10⁷ Nm/rad to 2.70×10¹⁰ Nm/rad. The tilt stiffness of the fanshaft mounting structure may be in the range from 2.1×10⁷ Nm/rad to1×10¹⁰ Nm/rad.

According to a nineteenth aspect, there is provided a gas turbine enginefor an aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades; a gearbox that receives an input from the core shaft andoutputs drive to a fan shaft so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier arranged to have the plurality of planet gearsmounted thereon; and

-   -   a fan shaft mounting structure arranged to mount the fan shaft        within the engine, the fan shaft mounting structure comprising        at least two supporting bearings connected to the fan shaft, and        wherein:    -   the torsional stiffness of the planet carrier is greater than or        equal to 1.60×10⁸ Nm/rad; and    -   the tilt stiffness of the fan shaft mounting structure is        greater than or equal to 1.50×10⁷ Nm/rad.

The torsional stiffness of the planet carrier may be greater than orequal to 2.7×10⁸ Nm/rad. The torsional stiffness of the planet carriermay be in the range from 1.60×10⁸ Nm/rad to 1.00×10¹¹ Nm/rad. Thetorsional stiffness of the planet carrier may be in the range from2.7×10⁸ Nm/rad to 1×10¹⁰ Nm/rad.

The tilt stiffness of the fan shaft mounting structure may be greaterthan or equal to 2.1×10⁷ Nm/rad. The tilt stiffness of the fan shaftmounting structure may be in the range from 1.50×10⁷ Nm/rad to 2.70×10¹⁰Nm/rad. The tilt stiffness of the fan shaft mounting structure may be inthe range from 2.1×10⁷ Nm/rad to 1×10¹⁰ Nm/rad.

The fan may have a fan diameter in the range from 240 cm to 280 cm. Insuch an embodiment, the torsional stiffness of the planet carrier may begreater than or equal to 1.8×10⁸ Nm/rad or in the range from 1.8×10⁸Nm/rad to 4.8×10⁹ Nm/rad. Additionally, or alternatively, in such anembodiment, the tilt stiffness of the fan shaft mounting structure maybe greater than or equal to 2.1×10⁷ Nm/rad or in the range from 2.1×10⁷Nm/rad to 1.9×10¹⁰ Nm/rad.

Alternatively, the fan may have a fan diameter in the range from 330 cmto 380 cm. In such an embodiment, the torsional stiffness of the planetcarrier may be greater than or equal to 6.0×10⁸ Nm/rad or in the rangefrom 6.0×10⁸ Nm/rad to 2.2×10¹⁰ Nm/rad. Additionally or alternatively,in such an embodiment, the tilt stiffness of the fan shaft mountingstructure may be greater than or equal to 3.8×10⁷ Nm/rad or in the rangefrom 3.8×10⁷ Nm/rad to 2.7×10¹⁰ Nm/rad.

Any one or more of the following may apply to either of both of theprevious two aspects (e.g. the eighteenth and/or nineteenth aspects):

A first planet carrier stiffness ratio of:

$\frac{{the}{effective}{linear}{torsional}{stiffness}{of}{the}{planet}{carrier}}{{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

may be greater than or equal to 7.0×10⁻³. The first planet carrierstiffness ratio may be greater than or equal to 7.0×10⁻². The firstplanet carrier stiffness ratio may be in the range from 7.0×10⁻³ to1.9×10³. The first planet carrier stiffness ratio may be in the rangefrom 7.0×10⁻² to 9.0×10¹.

A first planet carrier stiffness product of:

(the effective linear torsional stiffness of the planet carrier)×(radialbending stiffness of the fan shaft mounting structure)

may be greater than or equal to 2.9×10¹⁸ (N/m)². The first planetcarrier stiffness product may be greater than or equal to 5.0×10¹⁸(N/m)². The first planet carrier stiffness product may be in the rangefrom 2.9×10¹⁸ (N/m)² to 8.0×10²² (N/m)². The first planet carrierstiffness product may be in the range from 5.0×10¹⁸ (N/m)² to 8.0×10²¹(N/m)².

A second planet carrier stiffness ratio defined as:

$\frac{{the}{torsional}{stiffness}{of}{the}{planet}{carrier}}{{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

may be greater than or equal to 6.0×10⁻³. The second planet carrierstiffness ratio may be greater than or equal to 6.0×10⁻². The secondplanet carrier stiffness ratio may be in the range from 6.0×10⁻³ to7.0×10³. The second planet carrier stiffness ratio may be in the rangefrom 6.0×10⁻² to 7.0×10².

A second planet carrier stiffness product is defined as:

(the torsional stiffness of the planet carrier)×(tilt stiffness of thefan shaft mounting structure),

and may be greater than or equal to 2.4×10¹⁵ (Nm/rad)². The secondplanet carrier stiffness product may be greater than or equal to 4.9×10⁵(Nm/rad)². The second planet carrier stiffness product may be in therange from 2.4×10¹⁵ (Nm/rad)² to 2.7×10²¹ (Nm/rad)². The second planetcarrier stiffness product may be in the range from 4.9×10¹⁵ (Nm/rad)² to2.0×10²⁰ (Nm/rad)².

The power transmitted by the gearbox may be greater than or equal to2.25×10⁷ W. The power transmitted by the gearbox may be greater than orequal to 2.5×10⁷ W. The power transmitted by the gearbox may be in therange from 2.25×10⁷ W to 1.00×10⁸ W. The power transmitted by thegearbox may be in the range from 2.5×10⁷ W to 8.0×10⁷ W.

The moment of inertia of the fan may be greater than or equal to7.40×10⁷ kgm². The moment of inertia of the fan may be greater than orequal to 8.3×10⁷ kgm². The moment of inertia of the fan may be in therange from 7.40×10⁷ kgm² to 9.00×10⁸ kgm². The moment of inertia of thefan may be in the range from 8.3×10⁷ kgm² to 6.5×10⁸ kgm².

The at least two supporting bearings may comprise a first supportingbearing and second supporting bearing. Both of the first and the secondsupporting bearings may be located at positions forward of the gearbox.The first supporting bearing may be located at a position forward of thegearbox and the second supporting bearing may be located at a positionrearward of the gearbox.

The fan shaft mounting structure may further comprise a third supportingbearing. The third supporting bearing may be located between the fan andthe gearbox.

The fan shaft may comprise a gearbox output shaft forming a relativelyflexible portion of the fan shaft, and the fan shaft mounting structuremay comprise a gearbox output shaft support structure having at leastone gearbox output shaft bearing arranged to support the gearbox outputshaft.

The fan shaft mounting structure may further comprise one or morenon-supporting softly mounted bearings.

Any one or more of the bearings provided as part of the fan shaftmounting structure may be double bearings.

According to an twentieth aspect, there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier arranged to have the plurality of planet gearsmounted thereon; and a fan shaft mounting structure arranged to mountthe fan shaft within the propulsor, the fan shaft mounting structurecomprising at least two supporting bearings connected to the fan shaft,and wherein:

-   -   the torsional stiffness of the planet carrier is greater than or        equal to 1.6×10⁸ Nm/rad and the radial bending stiffness of the        fan shaft mounting structure is greater than or equal to        7.00×10⁸ N/m.

The propulsor of the twentieth aspect may have some or all of thefeatures described above with respect to the gas turbine engine of theeighteenth aspect, and may be a gas turbine engine in some embodiments.

According to an twenty-first aspect, there is provided a propulsor foran aircraft, comprising: a fan comprising a plurality of fan blades; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier arranged to have the plurality of planet gearsmounted thereon; and a fan shaft mounting structure arranged to mountthe fan shaft within the propulsor, the fan shaft mounting structurecomprising at least two supporting bearings connected to the fan shaft,and wherein:

-   -   the torsional stiffness of the planet carrier is greater than or        equal to 1.6×10⁸ Nm/rad and the tilt stiffness of the fan shaft        mounting structure is greater than or equal to 1.50×10⁷ Nm/rad.

The propulsor of the twenty-first aspect may have some or all of thefeatures described above with respect to the gas turbine engine of thenineteenth aspect, and may be a gas turbine engine in some embodiments.

The features of the twentieth and twenty-first aspects may be combined.According to twenty-second aspect, there is provided a propulsor for anaircraft, comprising: a fan comprising a plurality of fan blades; agearbox; a power unit for driving the fan via the gearbox, wherein thegearbox is arranged to receive an input from the power unit via a coreshaft and output drive to a fan shaft so as to drive the fan at a lowerrotational speed than the core shaft, the gearbox being an epicyclicgearbox comprising a sun gear, a plurality of planet gears, a ring gear,and a planet carrier arranged to have the plurality of planet gearsmounted thereon; and a fan shaft mounting structure arranged to mountthe fan shaft within the propulsor, the fan shaft mounting structurecomprising at least two supporting bearings connected to the fan shaft,and wherein:

-   -   a) the torsional stiffness of the planet carrier is greater than        or equal to 1.6×10⁸ Nm/rad and the radial bending stiffness of        the fan shaft mounting structure is greater than or equal to        7.00×10⁸ N/m; and/or    -   b) the torsional stiffness of the planet carrier is greater than        or equal to 1.6×10⁸ Nm/rad and the tilt stiffness of the fan        shaft mounting structure is greater than or equal to 1.50×10⁷        Nm/rad.

The propulsor of the twenty-second aspect may have some or all of thefeatures described above with respect to the gas turbine engine of theeighteenth or nineteenth aspects, and may be a gas turbine engine insome embodiments.

According to a twenty-third aspect, there is provided a method ofoperating a gas turbine engine for an aircraft comprising an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; a gearbox that receives aninput from the core shaft and outputs drive to a fan shaft so as todrive the fan at a lower rotational speed than the core shaft, thegearbox being an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier arranged to have theplurality of planet gears mounted thereon; and a fan shaft mountingstructure arranged to mount the fan shaft within the engine, the fanshaft mounting structure comprising at least two supporting bearingsconnected to the fan shaft, and wherein:

-   -   the torsional stiffness of the planet carrier is greater than or        equal to 1.6×10⁸ Nm/rad and the radial bending stiffness of the        fan shaft mounting structure is greater than or equal to        7.00×10⁸ N/m,        the method comprising operating the gas turbine engine to        provide propulsion for the aircraft under cruise conditions.

The method of the twenty-third aspect may be a method of operating thegas turbine engine or the propulsor of the eighteenth aspect ortwentieth aspect respectively. Any of the features, ratios andparameters introduced above in connection with the eighteenth ortwentieth aspect also therefore apply to the twenty-third aspect.

According to a twenty-fourth aspect, there is provided a method ofoperating a gas turbine engine for an aircraft comprising an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades; a gearbox that receives aninput from the core shaft and outputs drive to a fan shaft so as todrive the fan at a lower rotational speed than the core shaft, thegearbox being an epicyclic gearbox comprising a sun gear, a plurality ofplanet gears, a ring gear, and a planet carrier arranged to have theplurality of planet gears mounted thereon; and a fan shaft mountingstructure arranged to mount the fan shaft within the engine, the fanshaft mounting structure comprising at least two supporting bearingsconnected to the fan shaft, and wherein:

-   -   the torsional stiffness of the planet carrier is greater than or        equal to 1.6×10⁸ Nm/rad and the tilt stiffness of the fan shaft        mounting structure is greater than or equal to 1.50×10⁷ Nm/rad,    -   the method comprising operating the gas turbine engine to        provide propulsion for the aircraft under cruise conditions.

The method of the twenty-fourth aspect may be a method of operating thegas turbine engine or the propulsor of the nineteenth aspect ortwenty-first aspect respectively. Any of the features, ratios andparameters introduced above in connection with the nineteenth ortwenty-first aspect also therefore apply to the twenty-fourth aspect.

The twenty-third and twenty fourth aspects may be combined. According toa twenty-fifth aspect, there is provided a method of operating a gasturbine engine for an aircraft comprising an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades; a gearbox that receives an inputfrom the core shaft and outputs drive to a fan shaft so as to drive thefan at a lower rotational speed than the core shaft, the gearbox beingan epicyclic gearbox comprising a sun gear, a plurality of planet gears,a ring gear, and a planet carrier arranged to have the plurality ofplanet gears mounted thereon; and a fan shaft mounting structurearranged to mount the fan shaft within the engine, the fan shaftmounting structure comprising at least two supporting bearings connectedto the fan shaft, and wherein:

-   -   a) the torsional stiffness of the planet carrier is greater than        or equal to 1.6×10⁸ Nm/rad and the radial bending stiffness of        the fan shaft mounting structure is greater than or equal to        7.00×10⁸ N/m; and/or    -   b) the torsional stiffness of the planet carrier is greater than        or equal to 1.6×10⁸ Nm/rad and the tilt stiffness of the fan        shaft mounting structure is greater than or equal to 1.50×10⁷        Nm/rad,    -   the method comprising operating the gas turbine engine to        provide propulsion for the aircraft under cruise conditions.

The method of the twenty-fifth aspect may be a method of operating thegas turbine engine or the propulsor of the eighteenth, nineteenth, ortwenty-second aspect. Any of the features, ratios and parametersintroduced above in connection with the eighteenth, nineteenth, ortwenty-second aspect also therefore apply to the twenty-fourth aspect.

In the eighteenth and nineteenth aspects above the torsional stiffnessof the carrier may alternatively be defined as the effective lineartorsional stiffness of the carrier as defined elsewhere herein. The sameapplies to any of the twentieth to twenty-fifth aspects.

The inventor has discovered that designing the gas turbine engine sothat the torsional stiffness of the carrier and the radial bending ortilt stiffness of the fan shaft mounting structure are within thespecified ranges a high propulsive efficiency can be achieved. Theinventor has found that a stiffness of the fan shaft mounting structurewithin the specified range provides improved location of the fan so asto reduce any performance loss due to problems with fan tip clearancecontrol. The inventor has also found that by designing the gearbox sothat the torsional stiffness of the carrier is in the specified rangethe overall weight of the gearbox can be minimised so as to helpmaintain a low specific fuel consumption (SFC).

In other aspects, value ranges for the product or ratio of the torsionalstiffness of the carrier and the radial bending or tilt stiffness of thefan shaft mounting structure may be specified instead of, or as well as,absolute value ranges.

According to one such aspect, the eighteenth aspect introduced above maybe formulated as an aspect providing a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; a gearbox that receives an input from the core shaft and outputsdrive to a fan shaft so as to drive the fan at a lower rotational speedthan the core shaft, the gearbox being an epicyclic gearbox comprising asun gear, a plurality of planet gears, a ring gear, and a planet carrierarranged to have the plurality of planet gears mounted thereon; and

-   -   a fan shaft mounting structure arranged to mount the fan shaft        within the engine, the fan shaft mounting structure comprising        at least two supporting bearings connected to the fan shaft, and        wherein:        -   a) a first planet carrier stiffness ratio of:

$\frac{{the}{effective}{linear}{torsional}{stiffness}{of}{the}{planet}{carrier}}{{radial}{bending}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   -   is greater than or equal to 7.0×10⁻³, greater than or equal            to 7.0×10⁻², in the range from 7.0×10⁻³ to 1.9×10³, or in            the range from 7.0×10⁻² to 9.0×10¹; and/or        -   b) a first planet carrier stiffness product of:

(the effective linear torsional stiffness of the planet carrier)×(radialbending stiffness of the fan shaft mounting structure)

-   -   -   is greater than or equal to 2.9×10¹⁸ (N/m)², greater than or            equal to 5.0×10¹⁸ (N/m)², in the range from 2.9×10¹⁸ (N/m)²            to 8.0×10²² (N/m)², or in the range from 5.0×10¹⁸ (N/m)² to            8.0×10²¹ (N/m)².

According to another such aspect, the nineteenth aspect introduced abovemay be formulated as an aspect providing a gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; a gearbox that receives an input from the core shaft and outputsdrive to a fan shaft so as to drive the fan at a lower rotational speedthan the core shaft, the gearbox being an epicyclic gearbox comprising asun gear, a plurality of planet gears, a ring gear, and a planet carrierarranged to have the plurality of planet gears mounted thereon; and

-   -   a fan shaft mounting structure arranged to mount the fan shaft        within the engine, the fan shaft mounting structure comprising        at least two supporting bearings connected to the fan shaft, and        wherein:        -   a) a second planet carrier stiffness ratio is defined as:

$\frac{{the}{torsional}{stiffness}{of}{the}{planet}{carrier}}{{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}}$

-   -   -   is greater than or equal to 6.0×10⁻³, greater than or equal            to 6.0×10⁻², in the range from 6.0×10⁻³ to 7.0×10³, or in            the range from 6.0×10⁻² to 7.0×10²; and/or        -   b) a second planet carrier stiffness product is defined as:

(the torsional stiffness of the planet carrier)×(tilt stiffness of thefan shaft mounting structure),

-   -   and is greater than or equal to 2.4×10⁵ (Nm/rad)², greater than        or equal to 4.9×10⁵ (Nm/rad)², in the range from 2.4×10¹⁵        (Nm/rad)² to 2.7×10²¹ (Nm/rad)², or in the range from 4.9×10¹⁵        (Nm/rad)² to 2.0×10²⁰ (Nm/rad)².

The skilled person would appreciate that method and propulsor aspectsmay be formulated accordingly.

In any of the preceding aspects, any one or more of the following mayapply as applicable:

The turbine may be a first turbine, the compressor may be a firstcompressor, and the core shaft may be a first core shaft. The enginecore may further comprise a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

The gearbox may have a gear ratio in any range disclosed herein, forexample a gear ratio in the range from 3.2 to 4.5, and optionally from3.2 to 4.0.

The gas turbine engine may have a specific thrust in any range disclosedherein, for example a specific thrust in the range from 70 to 90 NKg⁻¹.

The gas turbine engine may have a bypass ratio at cruise conditions inany range disclosed herein, for example in the range from 12.5 to 18,and optionally from 13 to 16.

The fan may have a fan diameter greater than 240 cm and less than orequal to 380 cm. The fan may have a fan diameter greater than 300 cm andless than or equal to 380 cm. The fan may have a fan diameter in therange from 240 cm to 280 cm. The fan may have a fan diameter in therange from 330 cm to 380 cm.

The method of any of the aspects defined above may further comprisedriving the gearbox with an input torque of:

-   -   i) greater than or equal to 10,000 Nm, and optionally of 10,000        to 50,000 Nm at cruise; and/or    -   ii) greater than or equal to 28,000 Nm, and optionally of 28,000        to 135,000 Nm at max-take off conditions.

For any parameter or ratio of parameters X claimed or disclosed herein,a limit on the values that X can take that is expressed as “X is greaterthan or equal to Y” can alternatively be expressed as “1/X is less thanor equal to 1/Y”. Any of the ratios or parameters defined in the aspectsand statements above may therefore be expressed as “1/X is less than orequal to 1/Y” rather than “X is greater than or equal to Y”. In suchcases, zero can be taken as a lower bound.

Various parameters of the gearbox and its mounting structure, and/or ofthe engine more generally, may be adjusted to allow the engine to meetthe specifications of the various aspects summarised above. Comments onvarious such parameters are provided below.

The inventor has discovered that decreasing the stiffness (radialbending and/or tilt stiffness) of the fan shaft mounting structureoutside of the ranges defined herein would result in undesirablevibrations at low modal frequencies (the skilled person would appreciatethat the lower modal vibrations have larger amplitudes/deflections thanthe higher modes, and so are more important to avoid). This may be afunction of the size of the gearbox and its configuration.

The inventor has also found that increasing the fan shaft mountingstructure radial bending/tilt stiffness above the ranges defined hereinwould lead to excessive weight increase with little practicalperformance benefit. The inventor has appreciated that the maximumstiffness will be affected by the engineering limit of the material fromwhich the fan shaft mounting structure is made. The materials from whichthe fan shaft mounting structure is made (often steels) may, forexample, have a Young's modulus in the range from 100 to 250 GPa, or 105to 215 GPa, and optionally around 210 GPa—different grades of steel, orother types of metal, may be selected to achieve different stiffnessesfor the same size and geometry. For example, steels with a Young'smodulus in the range 190 to 215 GPa, titanium alloys with a Young'smodulus in the range 105 to 120 GPa, or a metal such as titanium with aYoung's modulus of around 110 GPa may be used in various embodiments.The inventor has discovered that increasing the stiffness beyond theranges defined herein using materials such as these would add excessiveweight while providing little or no practical performance benefit (e.g.in locating of the fan as described above).

The inventor has found that decreasing the radial bending and/or tiltstiffness of the fan shaft (at the input to the fan or the output of thegearbox) outside of the ranges defined herein would lead to undesirabledynamic effects such as lateral vibration. In particular, the minimumstiffness defined by the ranged specified herein allows vibrations atlow modal frequencies to be reduced or avoided (the skilled person wouldappreciate that the lower modal vibrations have largeramplitudes/deflections than the higher modes, and so are more importantto avoid). This may be a function of the size of the gearbox and itsconfiguration.

The inventor has also found that an upper limit of the fan shaft radialbending and/or tilt stiffness is affected by the fundamental propertiesof the material or materials from which it is made. For example, amaximum stiffness is affected by the engineering limit of the materialfrom which the fan shaft is made. The materials from which the fan shaftis made (often steels) may, for example, have a Young's modulus in therange from 100 to 250 GPa, or 105 to 215 GPa, and optionally around 210GPa—different grades of steel, or other types of metal, may be selectedto achieve different stiffnesses for the same size and geometry. Forexample, steels with a Young's modulus in the range 190 to 215 GPa,titanium alloys with a Young's modulus in the range 105 to 120 GPa, or ametal such as titanium with a Young's modulus of around 110 GPa may beused in various embodiments. The inventor has found increasing thestiffness outside of the ranges defined herein using materials such asthese would lead to excessive weight with no practical gain inperformance (e.g. no further practical gain in locating of the fan asdescribed above).

Regarding the first and second bearing separation distances, theinventor has discovered that increasing distance (d₁) outside of therange defined herein results in undesirable lateral vibrations caused bylow modal frequency vibrations and inadequate fan tip control. Theinventor has also found that decreasing d₁ outside of the range definedherein would result in design space problems e.g. making it difficultfor the gearbox to fit within the engine architecture. For example, theinventor has taken into account the need to fit other components withinthe engine. The inventor has found that the ranges for the distancesspecified herein provide a balance of these factors while giving thedesired benefits of gearbox isolation and location of the fan.

The inventor has found that increasing the distance between the gearboxoutput position and the closest supporting bearing forward of thegearbox (d₂) defined herein outside of the claimed range would notprovide suitable isolation for the gearbox and so would not avoidtransmitting damaging loads into it. As described elsewhere, this may bea function of the size and configuration of the gearbox. The inventorhas found that decreasing the distance below the specified range wouldlead to problems relating to the design space available so that thegearbox can fit in the engine.

Regarding the torsional stiffness of the carrier, the inventor has foundthat the upper limit of the ranges defined is affected by theengineering limits of typical materials chosen for the carrier, whichare often steels, and gearbox size. The materials of which the carrieris made (often steels) may, for example, have a Young's modulus in therange from 100 to 250 GPa, or 105 to 215 GPa, and optionally around 210GPa—different grades of steel, or other types of metal, may be selectedto achieve different stiffnesses for the same size and geometry. Forexample, steels with a Young's modulus in the range 190 to 215 GPa,titanium alloys with a Young's modulus in the range 105 to 120 GPa, or ametal such as titanium with a Young's modulus of around 110 GPa may beused in various embodiments. The inventor has found that increasing thetorsional stiffness beyond the ranges defined herein creates excessivecomponent weight for minimal improvement in operation.

The inventor has also found that for the torsional stiffness of thecarrier the lower limit of the ranges defined herein is affected by amaximum allowable deflection—the inventor has appreciated thatdisplacement may create mis-alignment in the gears and bearings and thata certain misalignment may be tolerated but that a larger displacementcould deleteriously affect running of the engine—a minimum stiffness maytherefore be selected to maintain displacement within acceptable bounds.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

The gas turbine engine may comprise a gearbox that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft. The input to thegearbox may be directly from the core shaft, or indirectly from the coreshaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed). The output from the gearbox may be directly to a fanshaft, or indirectly to the fan shaft, for example via a spur shaftand/or gear.

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform. The radius of the fan may be measuredbetween the engine centreline and the tip of a fan blade at its leadingedge. The fan diameter (which may simply be twice the radius of the fan)may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm,250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm(around 110 inches), 290 cm (around 115 inches), 300 cm (around 120inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches),340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm(around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165inches). The fan diameter may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 240 cm to 280cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. Thefan tip loading may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, a maximum take-off (MTO) condition has the conventionalmeaning. Maximum take-off conditions may be defined as operating theengine at International Standard Atmosphere (ISA) sea level pressure andtemperature conditions+15° C. at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditions forthe engine may therefore be defined as operating the engine at a maximumtake-off thrust (for example maximum throttle) for the engine at ISA sealevel pressure and temperature+15° C. with a fan inlet velocity of 0.25Mn.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

Whilst in the arrangements described herein the source of drive for thepropulsive fan is provided by a gas turbine engine, the skilled personwill appreciate the applicability of the gearbox configurationsdisclosed herein to other forms of aircraft propulsor comprisingalternative drive types. For example, the above-mentioned gearboxarrangements may be utilised in aircraft propulsors comprising apropulsive fan driven by an electric motor. In such circumstances, theelectric motor may be configured to operate at higher rotational speedsand thus may have a lower rotor diameter and may be more power-dense.The gearbox configurations of the aforesaid aspects may be employed toreduce the rotational input speed for the fan or propeller to allow itto operate in a more favourable efficiency regime. Thus, according to anaspect, there is provided an electric propulsion unit for an aircraft,comprising an electric machine configured to drive a propulsive fan viaa gearbox, the gearbox and/or its inputs/outputs/supports and/or thestructure by which the fan shaft driving the fan is supported being asdescribed and/or claimed herein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

As used herein, a range “from value X to value Y” or “between value Xand value Y”, or the like, denotes an inclusive range; including thebounding values of X and Y. As used herein, the term “axial plane”denotes a plane extending along the length of an engine, parallel to andcontaining an axial centreline of the engine, and the term “radialplane” denotes a plane extending perpendicular to the axial centrelineof the engine, so including all radial lines at the axial position ofthe radial plane. Axial planes may also be referred to as longitudinalplanes, as they extend along the length of the engine. A radial distanceor an axial distance is therefore a distance in a radial or axial plane,respectively.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic diagram illustrating the radial bending stiffnessof a cantilevered beam;

FIG. 5 is a schematic diagram illustrating the tilt stiffness of acantilevered beam;

FIG. 6 is a schematic diagram illustrating the torsional stiffness of ashaft;

FIG. 7 is a schematic close up sectional side view of a portion of a gasturbine engine around its gearbox;

FIGS. 8 and 9 are schematic diagrams illustrating the radial bendingstiffness a fan shaft mounting structure;

FIG. 10 is a close-up view showing only the bearings supporting the fanshaft to illustrate the measurement of the radial bending stiffness;

FIGS. 11 and 12 are schematic diagrams illustrating the tilt stiffnessthe fan shaft mounting structure;

FIG. 13 is a close-up view showing only the bearings supporting the fanshaft to illustrate the measurement of the tilt stiffness;

FIGS. 14 to 19 are schematic diagrams illustrating various embodimentsof the fan shaft mounting structure;

FIG. 20 is a schematic diagram illustrating the gearbox output positionfor a gearbox in a star configuration;

FIG. 21 is a schematic diagram illustrating the gearbox output positionfor a gearbox in a planetary configuration;

FIGS. 22 and 23 are schematic diagrams illustrating fan shaft end radialbending stiffnesses;

FIGS. 24 and 25 are schematic diagrams illustrating fan shaft end tiltstiffnesses;

FIG. 26 is a schematic diagram illustrating an alternative interfacebetween a fan shaft and fan;

FIG. 27 is a schematic diagram illustrating a planet carrier;

FIG. 28 is a schematic diagram illustrating the torsional stiffness ofthe carrier in side view;

FIG. 29 is a schematic diagram illustrating torsional stiffness of analternative carrier in front view;

FIG. 30 is a schematic diagram illustrating torsional stiffness of thecarrier of FIG. 29;

FIG. 31 is a schematic diagram illustrating a front view of a carriercomprising lugs;

FIG. 32 shows an aircraft having a gas turbine engine attached to eachwing;

FIG. 33 shows a method of operating a gas turbine engine on an aircraft;and

FIG. 34 shows a graph of applied load against displacement to illustratemeasurement of the stiffness of a component.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

The linkages 36 may be referred to as a fan shaft 36, the fan shaft 36optionally comprising two or more shaft portions coupled together. Forexample, the fan shaft 36 may comprise a gearbox output shaft portion 36a extending from the gearbox 30 and a fan portion 36 b extending betweenthe gearbox output shaft portion and the fan 23. In the embodiment shownin FIGS. 1 and 2, the gearbox 30 is a planetary gearbox and the gearboxoutput shaft portion 36 a is connected to the planet carrier 34—it maytherefore be referred to as a carrier output shaft 36 a. In stargearboxes 30, the gearbox output shaft portion 36 a may be connected tothe ring gear 38—it may therefore be referred to as a ring output shaft36 a. In the embodiment shown in FIGS. 1 and 2, the fan portion 36 b ofthe fan shaft 36 connects the gearbox output shaft portion 36 a to thefan 23. The output of the gearbox 30 is therefore transferred to the fan23, to rotate the fan, via the fan shaft 36. In alternative embodiments,the fan shaft 36 may comprise a single component, or more than twocomponents. Unless otherwise indicated or apparent to the skilledperson, anything described with respect to an engine 10 with a stargearbox 30 may equally be applied to an engine with a planetary gearbox30, and vice versa.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate. In various other exemplaryembodiments the gearbox may be any other type of gearbox, and so may notbe an epicyclic gearbox.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility as defined orclaimed herein. By way of further example, any suitable arrangement ofthe bearings between rotating and stationary parts of the engine (forexample between the input and output shafts from the gearbox and thefixed structures, such as the gearbox casing) may be used, and thedisclosure is not limited to the exemplary arrangement of FIG. 2. Forexample, where the gearbox 30 has a star arrangement (described above),the skilled person would readily understand that the arrangement ofoutput and support linkages and bearing locations would typically bedifferent to that shown by way of example in FIG. 2 (for example asdescribed in connection with other embodiments disclosed herein whichhave a star gearbox arrangement).

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

The following general definitions of stiffnesses may be used herein:

Radial Bending Stiffness

A radial bending stiffness is a measure of deformation for a given forceapplied in any one selected radial direction (i.e. any directionperpendicular to and passing through the engine axis). The radialbending stiffness is defined with reference to FIG. 4 in terms of thedeformation of a cantilevered beam 401. As illustrated in FIG. 4, aforce, F, applied at the free end of the beam in a directionperpendicular to the longitudinal axis of the beam causes a linearperpendicular deformation, δ. The radial bending stiffness is the forceapplied for a given linear deformation i.e. F/δ. In the presentapplication, the radial bending stiffness is taken relative to therotational axis of the engine 9, and so relates to the resistance tolinear deformation in a radial direction of the engine caused by aradial force. The beam, or equivalent cantilevered component, extendsalong the axis of rotation of the engine, the force, F, is appliedperpendicular to the axis of rotation of the engine, along any radialdirection, and the displacement 6 is measured perpendicular to the axisof rotation, along the line of action of the force. The radial bendingstiffness as defined herein has SI units of N/m. In the presentapplication, unless otherwise stated, the radial bending stiffness istaken to be a free-body stiffness i.e. stiffness measured for acomponent in isolation in a cantilever configuration, without othercomponents present which may affect its stiffness.

When the force is applied perpendicular to the cantilevered beam, and atthe free end of the beam, the resultant curvature is not constant butrather increases towards the fixed end of the beam.

Tilt Stiffness

A tilt stiffness is defined with reference to FIG. 5, which shows theresulting deformation of a cantilevered beam 401 under a moment Mapplied at its free end. The tilt stiffness is a measure of theresistance to rotation of a point on the component at which a moment isapplied. As can be seen in FIG. 5, an applied moment at the free end ofthe cantilevered beam induces a constant curvature along the length ofthe beam between its free and fixed ends. The applied moment M causes arotation θ of the point at which it is applied. For any component ofconstant section (like the beam), the angle θ is constant along thelength of the component. The tilt stiffness as defined herein thereforehas SI units of Nm/rad.

The tilt stiffness may be expressed as an effective linear tiltstiffness for a component having a given radius by expressing the tiltstiffness in terms of a pair of equal and opposite forces, F, acting ateither end of that radius (rather than the moment) and the arcdisplacement at that radius (i.e. displacement measured along acircumference of a circle having that radius). An approximate or overalltilt angle, a, may be defined for the purposes of calculating theeffective linear stiffness. The arc displacement may be referred to asrα. The effective linear tilt stiffness is given by the ratio ofeffective force divided by the displacement, F/rα and has the units N/m.

Torsional Stiffness

Torsional stiffness is a measure of deformation for a given torque. FIG.6 illustrates the definition of the torsional stiffness of a shaft 401or other body. A torque, τ, applied to the free end of the beam causes arotational deformation, θ (e.g. twist) along the length of the beam. Thetorsional stiffness is the torque applied for a given angle of twisti.e. τ/θ. The torsional stiffness has SI units of Nm/rad.

An effective linear torsional stiffness may be determined for acomponent having a given radius. The effective linear torsionalstiffness is defined in terms of an equivalent tangential force appliedat a point on that radius (with magnitude of torque divided by theradius) and the distance δ (with magnitude of the radius multiplied byθ) moved by a point corresponding to the rotational deformation θ of thecomponent.

The following general definitions of other parameters may also be usedherein:

Torque

Torque, which may also be referred to as moment, is the rotationalequivalent of linear force, and can be thought of as a twist to anobject.

The magnitude, τ, of torque, τ, of a body depends on three quantities:the force applied (F), the lever arm vector connecting the origin to thepoint of force application (r), and the angle (A) between the force andlever arm vectors:

τ=r×F

τ=|τ|=r×F|=|r∥F|sin A

where

-   -   τ is the torque vector and τ is the magnitude of the torque;    -   r is the position vector or “lever arm” vector (a vector from        the selected point on the body to the point where the force is        applied);    -   F is the force vector;    -   × denotes the cross product; and    -   A is the angle between the force vector and the lever arm vector        (sin(A) is therefore one when the force vector is perpendicular        to the position vector, such that τ=rF, i.e. magnitude of the        force multiplied by distance between the selected point on the        body and the point of application of the force).

Torque has dimensions of [force]×[distance] and may be expressed inunits of Newton metres (N·m).

The net torque on a body determines the rate of change of the body'sangular momentum.

Moment of Inertia

Moment of inertia, otherwise known as angular mass or rotationalinertia, is a quantity that determines the torque needed for a desiredangular acceleration of a body about a rotational axis—this issubstantially equivalent to how mass determines the force needed for aparticular acceleration.

Moment of inertia depends on the body's mass distribution and the axischosen, with larger moments requiring more torque to change the body'srotation rate. Moment of inertia has dimensions of [mass]×[distance]²and may be expressed in units of kilogram meter squared (kg. m²).

Moment of inertia I is defined as the ratio of the net angular momentumL of a body to its angular velocity ω around a principal axis:

$I = \frac{L}{\omega}$

Provided that the shape of the body does not change, its moment ofinertia appears in Newton's law of motion as the ratio of an appliedtorque τ on a body to the angular acceleration α around the principalaxis:

τ=Iα

For bodies constrained to rotate in a plane, only the moment of inertiaabout an axis perpendicular to the plane matters, and I can therefore berepresented as a scalar value. The skilled person would appreciate thata fan of a gas turbine engine (and more generally a fan rotor of the gasturbine engine comprising the fan disc and blades, and optionally alsothe fan shaft and/or other related components) is constrained to rotatein only one plane—a plane perpendicular to the engine axis—and that thefan's moment of inertia can therefore be defined by a single, scaler,value.

The fan's moment of inertia about the engine axis can therefore bemeasured or defined using any standard methodology.

More specific definitions of stiffnesses and other parameters relatingto embodiments described herein are provided below for ease ofunderstanding.

Fan Shaft Mounting Structure Stiffness

An embodiment of the gas turbine engine having a star configurationgearbox is shown in FIG. 7. Corresponding reference numbers to thoseused when describing the embodiment of FIGS. 1 and 2 have been used.FIG. 7 illustrates a close up of the engine core that shows the mountingof the fan shaft 36. The fan shaft 36 is mounted within the engine by afan shaft mounting structure 503. The fan shaft mounting structure 503comprises at least two bearings connected to or otherwise in engagementwith the fan shaft at points spaced apart axially along the length ofthe engine. The fan shaft mounting structure 503 may take a number ofdifferent forms, and may comprise one or more separate supportingstructures provided to support the fan shaft. It may also include otherstructures provided to support the fan shaft such as inter-shaftbearings. It therefore includes any supporting structure that extendsbetween a bearing in contact with the fan shaft and a stationarystructure of the engine (e.g. of the engine core).

In the arrangement shown in FIG. 7, the fan shaft mounting structure 503comprises two bearings, a first supporting bearing 506 a and a secondsupporting bearing 506 b, via which it is coupled to the fan shaft 36.The supporting bearings 506 a, 506 a are spaced apart along the axiallength of the fan shaft 36. In the described arrangement, bothsupporting bearings 506 a, 506 b are provided at positions that areforward of the gearbox 30. In other arrangements, one of the twosupporting bearing 506 a, 506 b used to support the fan shaft 36 may belocated at a position rearward of the gearbox 30, as will be describedlater. In yet other arrangements, more than two supporting bearings maybe provided as part of the fan shaft mounting structure or fan shaftmounting structure.

In addition to the supporting bearings 506 a, 506 b described above thefan shaft mounting structure may also comprise additional non-supportingbearings. These may be additional softly mounted bearing or provided aspart of a gearbox output shaft supporting structure as described in moredetail later. The supporting bearings may be defined as those having aprimary function of locating the fan shaft within the engine, ratherthan having a primary function of aligning other components such as thegearbox components.

The supporting bearings may be considered to be those transmitting anequal share of the total load that is an order of magnitude greater thanany non-supporting bearing. More specifically, a supporting bearing maybe defined as any bearing that transmits a load greater than 1/(10n) ofthe total load transmitted by the mounting structure of which it is apart, where n is the total number of bearings provided in that mountingstructure. For example, for a mounting structure having three bearings,any that contribute less than 1/30th (i.e. (⅓)10) would be consideredinsignificant and so not considered to be supporting bearings within themeaning of this application.

FIG. 7 shows a schematic example of a suitable fan shaft mountingstructure 503. Other forms of mounting structure may however be used tosupport the fan shaft. The fan shaft mounting structure 503 is coupledto the stationary support structure 24 of the engine so as to provide astationary mounting for the fan shaft within the engine (with rotationrelative to the static structure of the engine core provided by thebearings 506 a, 506 b). In the presently described arrangement, thestationary support structure 24 is an engine section stator (ESS) thatacts as both a structural component to provide a stationary mounting forcomponents such as the fan shaft 36, and as a guide vane provided todirect airflow from the fan 23. In other embodiments, the stationarysupporting structure 24 may comprise a strut extending across the coregas flow path and a separate stator vane provided to direct airflow, orany other suitable stationary structure relative to which the fan shaftmay be mounted.

The fan shaft mounting structure 503 is considered to comprise thecomponent or components extending between the point of contact betweeneach of the bearings 506 a, 506 b and the fan shaft 36 and thestationary support structure 24. Any number of separate components maybe provided between these points in order to provide a coupling betweenthe fan shaft 36 and the stationary support structure 24. The fan shaftmounting structure 503 is shown schematically in FIG. 7 for illustrationonly and other shapes and arrangements may be provided. For example, asdiscussed above additional bearings may be provided. These may benon-supporting softly mounted bearings, and/or additional bearingsprovided for redundancy in cases of high load transmission by thesupport structure. These additional bearings may be provided forward orrearward of each of the first and second bearings 506 a, 560 b. Anyadditional bearings provided for redundancy may protect against failuremodes of the primary components, either in response to normal operatingloads (e.g. fatigue) or higher loads from failure cases (e.g. Fan BladeOff). In some embodiments, any of the bearings (e.g. the first and/orsecond bearings 506 a, 506 b) may be double bearings formed by providinga pair of bearings in the same bearing housing. This arrangement may beused, for example, when the magnitude of the normal operating load wouldexceed that which would give reliable service with the use of singlebearings.

The fan shaft mounting structure 503 has a degree of flexibilitycharacterized by its radial stiffness and its tilt stiffness.

Fan Shaft Mounting Structure Radial Bending Stiffness:

The radial bending stiffness of the fan shaft mounting structure 503 isdefined with reference to FIGS. 8 and 9. The radial bending stiffnessrepresents the resistance of the fan shaft mounting structure 503 to aradial force. The radial bending stiffness is determined by treating thecomponents forming the fan shaft mounting structure 503 as a free bodythat is fixed at its point (or points) of connection with the stationarysupport structure (or structures) of the engine core (e.g. ESS 24). Inorder to evaluate the radial bending stiffness of the fan shaft mountingstructure 503 a radial force F is applied to a point on the fan (e.g. atthe interface between the fan blades and the fan hub) along its axialcentreline Z (i.e. along an axial centreline of the blades forming thefan). Specifically, the radial force F is applied along the fan axialcentreline (wherein the fan axial centreline is defined as the axialmidpoint of the fan blades, e.g. the midpoint of the fan blades in anaxial plane of the engine).

Application of this force causes a radial displacement 6 of the point ofcontact between the supporting bearings 506 a, 506 b and the fan shaft.Deformation of the fan shaft mounting structure 503 caused by theapplied force is illustrated in FIG. 9, with the shape when no force isapplied shown in broken lines for comparison. The force, F, is shownradially away from the engine axis 9, but could equivalently be a forcein a radial direction toward the centreline 9.

The radial bending stiffness of the fan shaft mounting structure 503 isdefined as the force F divided by the average displacement at thesupporting bearings provided as part of the fan shaft mounting structureforward of the gearbox 30. These displacements are illustrated in theclose-up view of FIG. 10 in which only the bearings are shown for easeof explanation. For the arrangement shown in FIGS. 9 and 10, the radialbending stiffness is therefore given by:

$\frac{F}{1/2( {\delta_{a} + \delta_{b}} )}$

The radial bending stiffness of the fan shaft mounting structure hasunits of N/m.

In various embodiments, the radial bending stiffness of the fan shaftmounting structure may be greater than or equal to 7.00×10⁸ N/m andoptionally greater than or equal to 1.25×10⁹ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the radial bendingstiffness of the fan shaft mounting structure may be greater than orequal to 7.0×10⁸ N/m. In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the radialbending stiffness of the fan shaft mounting structure may be greaterthan or equal to 1.4×10⁹ N/m.

In various embodiments, the radial bending stiffness of the fan shaftmounting structure may be in the range from 7.00×10⁸ N/m to 6.00×10¹¹N/m, and optionally in the range from 1.25×10⁹ N/m to 2.0×10¹¹ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the radial bendingstiffness of the fan shaft mounting structure may be in the range from7.0×10⁸ N/m to 5.0×10¹¹ N/m and optionally in the range from 7.0×10⁸ N/mto 2.3×10⁹ N/m (and may be equal to 1.5×10⁹ N/m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the radial bendingstiffness of the fan shaft mounting structure may be in the range from1.4×10⁹ N/m to 6.0×10¹¹ N/m, and optionally in the range from 1.4×10⁹N/m to 3.0×10⁹ N/m (and may be equal to 2.2×10⁹ N/m).

Fan Shaft Mounting Structure Tilt Stiffness:

The tilt stiffness of the fan shaft mounting structure 503 is determinedin a similar way to the radial bending stiffness except that a moment,M, is applied at a point on the fan blade axial centre line in place ofthe Force, F. An example of how the tilt stiffness can be determined isillustrated in FIGS. 11 and 12. The fan shaft mounting structure tiltstiffness represents its resistance to an applied moment. Referring toFIGS. 11 and 12, moment M is applied to the axial centre line of the fanblades (i.e. the same place as the force, F, in FIG. 8 described above).The tilt stiffness is determined by calculating a straight line shaftangle between a pair of bearing positions at which the fan shaft issupported. The application of the moment M causes displacements inopposite directions at two bearing positions as illustrated in FIG. 13,which shows a close-up view including only the bearings for ease ofexplanation. The straight line angle θ is defined as the change in angleof an axis R extending through the point of contact between the bearings506 a, 506 b and the fan shaft when the moment M is applied. As can beseen in FIG. 13, the angle θ extended between axis R when no moment isapplied and R′ when moment M is applied. The tilt stiffness is definedas M/θ, and has units of Nm/rad.

The tilt stiffness of the fan shaft mounting structure is defined bymeasuring the angle θ between the first two supporting bearings in adirection moving reward from the fan. In the arrangement shown in FIGS.11, 12 and 13 this corresponds to supporting bearings 506 a, 506 bprovided forward of the gearbox. In other arrangements the tiltstiffness may be defined using the displacement at a supporting bearingrearward of the gearbox (e.g. if only one supporting bearing is providedforward of the gearbox).

In various embodiments, the tilt stiffness of the fan shaft mountingstructure may be greater than or equal to 1.50×10⁷ Nm/rad, andoptionally greater than or equal to 2.1×10⁷ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the tilt stiffness of thefan shaft mounting structure may be greater than or equal to 2.1×10⁷Nm/rad or greater than or equal to 2.3×10⁷ Nm/rad. In some embodiments,for example in embodiments in which the fan diameter is in the rangefrom 330 to 380 cm, the tilt stiffness of the fan shaft mountingstructure may be greater than or equal to 3.8×10⁷ Nm/rad or greater thanor equal to 7.3×10⁷ Nm/rad.

In various embodiments, the tilt stiffness of the fan shaft mountingstructure may be in the range from 1.5×10⁷ Nm/rad to 2.70×10¹⁰ Nm/rad,and optionally in the range from 2.1×10⁷ Nm/rad to 1×10¹⁰ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the tilt stiffness of thefan shaft mounting structure may be in the range from 2.1×10⁷ Nm/rad to1.9×10¹⁰ Nm/rad and optionally in the range from 2.3×10⁷ Nm/rad to4.3×10⁷ Nm/rad (and may be equal to 3.3×10⁷ Nm/rad).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the tilt stiffness of thefan shaft mounting structure may be in the range from 3.8×10⁷ Nm/rad to2.7×10¹⁰ Nm/rad, and optionally in the range from 7.1×10⁷ Nm/rad to9.1×10⁷ Nm/rad (and may be equal to 8.1×10⁷ Nm/rad).

Equivalent radial and tilt stiffness may be defined for otherarrangements of fan shaft mounting structures. Various otherarrangements of fan shaft mounting structure 503 are shown in FIGS. 14to 19. These embodiments are provided as examples. The skilled personwill understand that others arrangements are possible and are consideredto fall within the scope of the present disclosure.

FIG. 14 shows a schematic representation of an arrangement in which thefan shaft 36 is supported by the fan shaft mounting structure alreadydescribed having first and second supporting bearings 506 a, 506 bforward of the gearbox 30. In all of FIGS. 14 to 19 the core shaft 26 issupported by two core shaft bearings 507 a, 507 b. These bearing areconnected to a stationary structure 24 a of the engine via a core shaftsupport structure. In the arrangement of FIG. 14, the fan shaft mountingstructure 503 comprises both the first and second bearings 506 a, 506 band the structure linking them to the stationary support structure 24(i.e. the fan shaft support structure 504). Any additional bearing orbearings provided as part of this structure is included in the fan shaftmounting structure 503. In the arrangement of FIG. 14, a non-supportingsoftly mounted bearing 506 a′ is also provided. As discussed above, theradial bending or tilt stiffness is determined without consideringdisplacements at this non-supporting bearing 506 a′.

FIG. 15 shows an arrangement in which the fan shaft is supported by afirst supporting bearing 506 a forward of the gearbox 30 and a secondsupporting bearing 506 b located at a position rearward of the gearbox30. In this embodiment, the fan shaft 36 extends through the gearbox 30and is mounted by supporting bearings 506 a, 506 b either side of thegearbox. The first supporting bearing 506 a is therefore located forwardof the gearbox 30 and the second supporting bearing 506 b is locatedrearward of the gearbox 30. In this embodiment, the second supportingbearing 506 b is an inter-shaft bearing between the fan shaft 36 and thecore shaft 26. For this arrangement the radial bending stiffness isdetermined by measuring the displacement at the first supporting bearing506 a only. The tilt stiffness is however determined by measuring thechange in angle of an axis linking the first and second supportingbearings 506 a, 506 b.

FIG. 16 shows an arrangement in which the fan shaft is supported by thefirst and second supporting bearings 506 a, 506 b either side of thegearbox. In this embodiment, the fan shaft mounting structure 503further comprises a third supporting bearing 506 c that is locatedbetween the first supporting bearing 506 a and the gearbox 30. In thisarrangement the radial bending stiffness of the fan shaft mountingstructure is determined by measuring an average displacement at thefirst and third supporting bearings 506 a, 506 c. Any displacement at ofthe second supporting bearing 506 b is not included as it is not forwardof the gearbox 30. The tilt stiffness is determined by measuring thechange in angle of an axis linking the first and third supportingbearings 506 a, 506 c as they are the first two bearings rearward of thefan.

FIG. 17 shows a modification to the arrangement of FIG. 14 in which thefan shaft 36 comprises a gearbox output shaft 36 a. The gearbox outputshaft 36 a forms a flexible portion of the fan shaft 36 (i.e. moreflexible relative to an intermediate portion, e.g. the fan portion 36 a,of the fan shaft to which it is connected) at the end at which itconnects to the gearbox 30. The gearbox output shaft 36 a is supportedby a bearing 508 a forming part of a gearbox output shaft supportstructure 41. In this embodiment, the fan shaft mounting structure 503therefore comprises both the first and second bearings 506 a, 506 b andthe structure linking them to the stationary support structure 24 (i.e.the fan shaft support structure 504) and the gearbox output shaftsupport structure 41 including its bearing 508 a. As the bearing 508 aof the gearbox output shaft support structure is a non-supportingbearing it is not taken into account when measuring the radial bendingstiffness. For this arrangement the radial bending stiffness is againdetermined by measuring the average of the displacement at the first andsecond supporting bearings 506 a, 506 b. The tilt stiffness is alsomeasured at the first and second bearings 506 a, 506 b.

FIG. 18 shows an arrangement in which the gearbox output shaft 36 aextends in an axial direction forward and rearwards of the gearbox 30,and is supported either side of the gearbox 30 by a first gearbox outputshaft bearing 508 a, and a second gearbox output shaft bearing 508 b.Both the first and second gearbox output shaft bearings 508 a, 508 bform part of the gearbox output shaft supporting structure 41 and so areconnected to a stationary structure of the engine. In this arrangement,the fan shaft mounting structure 503 comprises both the first and secondbearings 506 a, 506 b and the structure linking them to the stationarysupport structure 24 (i.e. the fan shaft support structure 504) and thegearbox output shaft support structure 41 including its two bearings.Again the output shaft bearings 508 a, 508 b are not included in themeasurement of the radial bending stiffness of the fan shaft mountingstructure as they are non-supporting bearings. The tilt stiffness isalso measured at the first and second supporting bearings.

FIG. 19 shows an arrangement having the same gearbox output shaftsupport as that of FIG. 16, but in which the second bearing 506 bprovided to support the fan shaft 36 is provided rearward of the gearbox30 as described in connection with FIG. 16. In this arrangement, the fanshaft mounting structure includes: the fan shaft support structure 504;the gearbox support structure 41 and its bearings; and the inter-shaftbearing 506 b and the components linking it to the stationary structureof the engine 24 a. The radial bending stiffness is determined bymeasuring a displacement of the first supporting bearing 506 a. Anydisplacement at the second supporting bearing 506 b is not included asit is not forward of the gearbox. Any displacement at the gearbox outputshaft bearings 508 a, 508 b is also not included as they arenon-supporting bearings. The tilt stiffness is determined by measuringthe angle of an axis linking the first supporting bearing 506 a and thesecond supporting bearing 506 b.

Fan Shaft Stiffness

The stiffness of the fan shaft is defined with reference to FIGS. 20 to26. The fan shaft 36 is defined as the torque transfer component thatextends from the output of the gearbox to the fan input. It thereforeincludes part or all of any gearbox output shaft and fan input shaftthat may be provided between those points. For the purposes of definingthe stiffness of the fan shaft 36 it is considered to extend between afan input position and a gearbox output position, and includes all ofthe torque transfer components between those points. It does nottherefore include any components of the gearbox itself (e.g. the planetcarrier or connecting plate coupled to it) which transmit discreteforces, rather than the fan shaft torque. The gearbox output positiontherefore may be defined as the point of connection between the fanshaft 36 and the gearbox 30. The fan input position may be defined asthe point of connection between the fan shaft 36 and the fan.

Referring to FIG. 20, where the gearbox is in a star configuration, thegearbox output position is defined as the point of connection 702between the ring gear 38 and the fan shaft 36. More specifically, it isthe point of connection to the annulus of the ring gear 38 (with anyconnection component extending from the outer surface of the annulusbeing considered to be part of the ring gear). Where the point ofconnection is formed by an interface extending in a direction having anaxial component, the point of connection is considered to be the axialcentreline of that interface as illustrated in FIG. 22.

The fan shaft 36 includes all torque transmitting components up to thepoint of connection 702 with the ring gear 38. It therefore includes anyflexible portions or linkages 704 making up the fan shaft 36 that may beprovided, and any connection(s) 706 (e.g. spline connections) betweenthem.

Where the gearbox 30 is in a planetary configuration, the gearbox outputposition is again defined as the point of connection between the fanshaft 36 and the gearbox 30. An example of this is illustrated in FIG.21, which shows a carrier comprising a forward plate 34 a and rearwardplate 34 b, with a plurality of pins 33 extending between them and onwhich the planet gears are mounted. The fan shaft 36 is connected to theforward plate 34 a via a connection 708 (e.g. a spline connection). Inan embodiment such as this, the gearbox output position is taken as anypoint on the interface between the fan shaft 36 and the forward plate 34a. The forward plate 34 a is considered to transmit discrete forces,rather than a single torque, and so is taken to be part of the gearbox30 rather than the fan shaft. FIG. 21 shows only one example of a typeof connection between the fan shaft and planet carrier 34. Inembodiments having different connection arrangements, the gearbox outputposition is still taken to be at the interface between componentstransmitting a torque (i.e. that are part of the fan shaft) and thosetransmitting discrete forces (e.g. that are part of the gearbox). Thespline connection 708 is only one example of a connection that may beformed between the fan shaft and gearbox (i.e. between the fan shaft andthe forward plate 34 b in the presently described arrangement). In otherembodiments, the interface which forms the gearbox output position maybe formed by, for example, a curvic connection, a bolted joint or othertoothed or mechanically dogged arrangement.

Referring to FIG. 22, the fan input position is defined as a point onthe fan shaft at the axial midpoint of the interface between the fan andthe fan shaft. In the presently described arrangement, the fan 23comprises a support arm 23 a arranged to connect the fan 23 to the fanshaft 36. The support arm 23 a is connected to the fan shaft by a splinecoupling 36 c (shown in FIG. 22) that extends along the length of aportion of the fan shaft 36. The fan input position is defined as theaxial midpoint of the spline coupling as indicated by axis Y in FIG. 22.The spline coupling shown in FIG. 22 is only one example of a couplingthat may form the interface between the fan and fan shaft. In otherembodiments, for example, a curvic connection, a bolted joint or othertoothed or mechanically dogged arrangement may be used.

FIG. 26 illustrates an arrangement in which an alternative coupling isprovided between the fan 23 and the fan shaft 36. Similarly to FIG. 22,the fan 23 is coupled to the fan shaft 36 via a support arm 23 a. Inthis arrangement however, a flange coupling 36 d is provided between thesupport arm 23 a and the fan shaft 36. In this embodiment, the supportarm 23 a is connected at the rear of the fan hub. The flange coupling 36d may be a curvic coupling. In other embodiments, other forms of flangecoupling may be provided. In the embodiment of FIG. 26, the fan inputposition is the axial midpoint of the flange coupling.

The fan shaft 36 has a degree of flexibility characterized by its radialbending stiffness and tilt stiffness.

Fan Shaft End Stiffness at Fan Input and Gearbox Output:

The stiffness of each end of the fan shaft where it couples to the fan23 and the gearbox 30 is defined with reference to FIGS. 22 to 25.

The radial bending stiffness of the fan shaft 36 at the input to the fan23 is measured by applying a force F₁ to the fan shaft at the fan shaftinput position defined above (illustrated in FIG. 22). The fan shaft 36is treated as a free body, and is held fixed at the position of all ofbearing positions at which it is supported i.e. the first and secondsupporting bearings 506 a,506 b in the embodiment of FIG. 22. As aresult of the application of the force F₁ the fan shaft 36 deforms sothat the fan shaft input position is displaced by a distance of δ₁ (seeFIG. 23). The radial bending stiffness of the fan shaft 36 at the inputto the fan 23 is then given by F₁/δ₁.

In various embodiments, the radial bending stiffness of the fan shaft atthe input to the fan may be greater than or equal to 3.00×10⁶ N/m, andoptionally greater than or equal to 6.3×10⁶ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the radial bendingstiffness of the fan shaft at the input to the fan may be greater thanor equal to 6.4×10⁶ N/m. In some embodiments, for example in embodimentsin which the fan diameter is in the range from 330 to 380 cm, the radialbending stiffness of the fan shaft at the input to the fan may begreater than or equal to 6.9×10⁶ N/m or greater than or equal to 8.9×10⁶N/m.

In various embodiments, the radial bending stiffness of the fan shaft atthe input to the fan may be in the range from 3.00×10⁶ N/m to 2.00×10⁹N/m, and optionally in the range from 6.3×10⁶ N/m to 1.0×10⁹ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the radial bendingstiffness of the fan shaft at the input to the fan may be in the rangefrom 6.4×10⁶ N/m to 1.0×10⁹ N/m and optionally in the range from 6.4×10⁶N/m to 7.6×10⁶ N/m (and may be equal to 7.0×10⁶ N/m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the radial bendingstiffness of the fan shaft at the input to the fan may be in the rangefrom 6.9×10⁶ N/m to 2.0×10⁹ N/m, and optionally in the range from8.9×10⁶ N/m to 1.1×10⁷ N/m (and may be equal to 9.9×10⁶ N/m).

The tilt stiffness of the fan shaft 36 at the input to the fan 23 ismeasured by applying a moment M₁ to the fan shaft at the fan shaft inputposition defined above (illustrated in FIG. 24). The fan shaft 36 isagain treated as a free body, and is held fixed at the position of allof the bearing positions at which it is supported i.e. the first andsecond supporting bearings 506 a,506 b in the arrangement of FIG. 24. Asa result of moment M₁ the fan shaft 36 deforms so that the fan shaftinput position is displaced by an angular displacement of θ₁ asillustrated in FIG. 25. The tilt stiffness of the fan shaft 36 at theinput to the fan 23 is then given by M₁/θ₁.

In various embodiments, the tilt stiffness of the fan shaft at the inputto the fan may be greater than or equal to 5.00×10⁵ Nm/rad, andoptionally greater than or equal to 9.0×10⁵ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the tilt stiffness of thefan shaft at the input to the fan may be greater than or equal to9.5×10⁵ Nm/rad. In some embodiments, for example in embodiments in whichthe fan diameter is in the range from 330 to 380 cm, the tilt stiffnessof the fan shaft at the input to the fan may be greater than or equal to1.5×10⁶ Nm/rad or greater than or equal to 2.5×10⁶ Nm/rad.

In various embodiments, the tilt stiffness of the fan shaft at the inputto the fan may be in the range from 5.00×10⁵ Nm/rad to 7.00×10⁸ Nm/rad,and optionally in the range from 9.0×10⁵ Nm/rad to 3.5×10⁸ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the tilt stiffness of thefan shaft at the input to the fan may be in the range from 9.5×10⁵Nm/rad to 3.5×10⁸ Nm/rad and optionally in the range from 9.5×10⁵ Nm/radto 1.9×10⁶ Nm/rad (and may be equal to 1.2×10⁶ Nm/rad).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the tilt stiffness of thefan shaft at the input to the fan may be in the range from 1.5×10⁶Nm/rad to 7.0×10⁸ Nm/rad, and optionally in the range from 2.5×10⁶Nm/rad to 4.5×10⁶ Nm/rad (and may be equal to 3.5×10⁶ Nm/rad).

The radial bending stiffness of the fan shaft 36 at the output of thegearbox 30 is measured by applying a force F₂ to the fan shaft at thegearbox output position defined above (illustrated in FIG. 22). The fanshaft 36 is treated as a free body, and is held fixed at the position ofall of bearing positions at which it is supported i.e. the first andsecond supporting bearings 506 a,506 b in the arrangement of FIG. 22. Asa result of force F₂ the fan shaft 36 deforms so that the gearbox outputposition is displaced by a distance of 62 (as illustrated in FIG. 23).The radial bending stiffness of the fan shaft 36 at the output of thegearbox is then given by F₂/δ₂.

In various embodiments, the radial bending stiffness of the fan shaft atthe output of the gearbox may be greater than or equal to 4.00×10⁶ N/m,and optionally greater than or equal to 3.7×10⁷ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the radial bendingstiffness of the fan shaft at the output of the gearbox may be greaterthan or equal to 3.7×10⁷ N/m. In some embodiments, for example inembodiments in which the fan diameter is in the range from 330 to 380cm, the radial bending stiffness of the fan shaft at the output of thegearbox may be greater than or equal to 3.9×10⁷ N/m or greater than orequal to 5.0×10⁷ N/m.

In various embodiments, the radial bending stiffness of the fan shaft atthe output of the gearbox may be in the range from 4.00×10⁶ N/m to1.5×10⁹ N/m, and optionally in the range from 3.7×10⁷ N/m to 1.0×10⁹N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the radial bendingstiffness of the fan shaft at the output of the gearbox may be in therange from 3.7×10⁷ N/m to 5.0×10⁸ N/m and optionally in the range from3.7×10⁷ N/m to 5.0×10⁷ N/m (and may be equal to 4.0×10⁷ N/m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the radial bendingstiffness of the fan shaft at the output of the gearbox may be in therange from 3.9×10⁷ N/m to 1.5×10⁹ N/m, and optionally in the range from5.0×10⁷ N/m to 9.0×10⁷ N/m (and may be equal to 7.0×10⁷ N/m).

The tilt stiffness of the fan shaft 36 at the output of the gearbox 30is measured by applying a moment M₂ to the fan shaft at the gearboxoutput position defined above (illustrated in FIG. 24). The fan shaft 36is again treated as a free body, and is held fixed at the position ofall of the bearing positions at which it is supported i.e. the first andsecond supporting bearings 506 a,506 b in the embodiment of FIG. 24. Asa result of moment M₂ the fan shaft 36 deforms so that the gearboxoutput position is displaced by an angular displacement of θ₂ asillustrated in FIG. 25. The tilt stiffness of the fan shaft 36 at theoutput of the gearbox is then given by M₂/θ₂.

In various embodiments, the tilt stiffness of the fan shaft at theoutput of the gearbox may be greater than or equal to 7.00×10⁴ Nm/rad,and optionally greater than or equal to 9.5×10⁵ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the tilt stiffness of thefan shaft at the output of the gearbox may be greater than or equal to9.5×10⁵. In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the tilt stiffness ofthe fan shaft at the output of the gearbox may be greater than or equalto 1.1×10⁶ Nm/rad and optionally may be greater than or equal to 2.6×10⁶Nm/rad.

In various embodiments, the tilt stiffness of the fan shaft at theoutput of the gearbox may be in the range from 7.00×10⁴ Nm/rad to7.00×10⁷ Nm/rad, and optionally in the range from 9.5×10⁵ Nm/rad to3.5×10⁷ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the tilt stiffness of thefan shaft at the output of the gearbox may be in the range from 9.5×10⁵Nm/rad to 2.0×10⁷ Nm/rad and optionally in the range from 9.5×10⁵ Nm/radto 2.2×10⁶ Nm/rad (and may be equal to 1.2×10⁶ Nm/rad).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the tilt stiffness of thefan shaft at the output of the gearbox may be in the range from 1.1×10⁶Nm/rad to 7.0×10⁷ Nm/rad, and optionally in the range from 2.6×10⁶Nm/rad to 4.6×10⁶ Nm/rad (and may be equal to 3.6×10⁶ Nm/rad).

Relative Distance to and Between Fan Shaft Supporting Bearings

Referring to FIG. 22, the fan shaft end stiffness is at least partlydetermined according to the respective distance between the fan inputposition, the gearbox output position and the bearings by which the fanshaft is supported. For example, the stiffness of the fan shaft at thefan input may depend on the axial distance d₁ between the fan inputposition and the closest supporting bearing rearward of the fan i.e. thefirst supporting bearing 506 a in the arrangement of FIG. 22. The axialposition of the first supporting bearing 506 a is taken to be its axialcentreline. The stiffness of the fan shaft at the gearbox output maydepend on the axial distance d₂ between the gearbox output position andthe closest supporting bearing forward of the gearbox i.e. the secondsupporting bearing 506 b in the arrangement of FIG. 22. The axialposition of the second supporting bearing 506 b is taken to be its axialcentreline. The closest supporting bearing forward of the gearbox outputposition does not include any non-supporting flexible or softly mountedbearings that also may be coupled to the fan shaft as described above inconnection with the measurement of the stiffness of the fan shaftmounting structure. It does not, for example, include the gearbox outputshaft bearing 508 a included in arrangements such as those shown inFIGS. 17, 18 and 19. By ‘supporting bearing’ we therefore again mean abearing transmitting a significant amount of the load transmitted by thefan shaft supporting structure during normal operation as definedelsewhere herein.

The gas turbine engine 10 may be configured such that the relativepositions of the bearings (e.g. axial distances d₁ and d₂) provide a fanshaft end stiffness within the desired range.

In various embodiments, distance d₁ may be greater than or equal to 0.12m, and optionally greater than or equal to 0.13 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₁ may be greaterthan or equal to 0.12 m or greater than or equal to 0.13 m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, distance d₁ may be greater than or equalto 0.13 m or greater than or equal to 0.15 m.

In various embodiments, distance d₁ may be in the range from 0.12 m to0.40 m, and optionally in the range from 0.13 m to 0.30 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₁ may be in therange from 0.12 m to 0.30 m and optionally in the range from 0.13 m to0.15 m (and may be equal to 0.14 m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, distance d₁ may be in therange from 0.13 m to 0.40 m and optionally in the range from 0.15 m to0.25 m (and may be equal to 0.20 m).

In various embodiments, distance d₂ may be greater than or equal to 0.15m, and optionally greater than or equal to 0.16 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₂ may be greaterthan or equal to 0.15 m or greater than or equal to 0.16 m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, distance d₂ may be greater than or equalto 0.16 m or greater than or equal to 0.20 m.

In various embodiments, distance d₂ may be in the range from 0.15 m to0.45 m, and optionally in the range from 0.16 m to 0.40 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₂ may be in therange from 0.15 m to 0.35 m and optionally in the range from 0.16 m to0.18 m (and may be equal to 0.17 m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, distance d₂ may be in therange from 0.16 m to 0.45 m and optionally in the range from 0.20 m to0.28 m (and may be equal to 0.24 m).

In the arrangement of FIG. 15, the distance d₁ corresponds to thedistance between the first supporting bearing 506 a and the fan inputposition. The distance d₂ corresponds to the distance between the firstsupporting bearing 506 a and the gearbox input position.

A bearing axial separation d₃ is defined as the axial distance betweenthe first supporting bearing 506 a and the second supporting bearing 506b as shown in FIG. 22. This distance therefore corresponds to the axialseparation between the closest supporting bearing rearward of the fan(i.e. the first supporting bearing 506 a) and the closest supportingbearing forward of the gearbox (i.e. the second supporting bearing 506b). Again, the axial position of the first and second supportingbearings 506 a,b is taken to be their respective axial centreline.

In various embodiments, distance d₃ may be in the range from 0.1 m to0.4 m, and optionally in the range from 0.18 m to 0.32 mm and furtheroptionally in the range from 0.20 m to 0.30 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₃ may be in therange from 0.19 m to 0.23 m (and may be equal to 0.21 m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, distance d₃ may be in therange from 0.26 m to 0.30 m (and may be equal to 0.28 m).

An axial distance d₄ is defined as the axial distance between the faninput position and the gearbox output position as can be seen in FIG. 22(e.g. between Y and X). The axial distance d₄ is equal to the sum ofdistances d₁, d₂ and d₃ defined above.

The axial distance d₃ may only be defined for arrangements in which thefirst and second supporting bearings 506 a, 506 b are located forward ofthe gearbox. The axial distance d₄ is however defined for allembodiments. Where only one supporting bearing is located forward of thegearbox distance d₃ is zero. An example of how d₁, d₂ and d₄ may bedefined for an arrangement in which a single supporting bearing isprovided forward of the gearbox is shown in FIG. 15.

In various embodiments, distance d₄ may be greater than or equal to 0.43m, and optionally greater than or equal to 0.46 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₄ may be greaterthan or equal to 0.43 m or greater than or equal to 0.48 m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, distance d₄ may be greater than or equalto 0.56 m or greater than or equal to 0.65 m.

In various embodiments, distance d₄ may be in the range from 0.43 m to0.95 m, and optionally in the range from 0.46 m to 0.85 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, distance d₄ may be in therange from 0.43 m to 0.62 m and optionally in the range from 0.48 m to0.56 m (and may be equal to 0.52 m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, distance d₄ may be in therange from 0.56 m to 0.95 m and optionally in the range from 0.65 m to0.76 m (and may be equal to 0.71 m).

Relative Fan and Gearbox Positions

Referring again to FIG. 7, a fan-gearbox axial distance 110 is definedas the axial distance between the output of the gearbox (i.e. the axialposition P of the gearbox output position) and the fan axial centrelineQ. The fan axial centreline is defined as the axial midpoint of the fanblades forming the fan (and may correspond to axis Z). The gearboxoutput position is defined as the point of connection between the fanshaft 36 and the gearbox as defined elsewhere herein. This may bedefined differently for different types of gearbox as describedelsewhere herein.

In various embodiments, the fan-gearbox axial distance may be greaterthan or equal to 0.35 m, and optionally greater than or equal to 0.37 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan-gearbox axialdistance may be greater than or equal to 0.38 m or greater than or equalto 0.40 m. In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the fan-gearbox axialdistance may be greater than or equal to 0.48 m or greater than or equalto 0.50 m.

In various embodiments, the fan-gearbox axial distance may be in therange from 0.35 m to 0.8 m, and optionally in the range from 0.37 m to0.75 m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan-gearbox axialdistance may be in the range from 0.38 m to 0.65 m and optionally in therange from 0.40 m to 0.44 m (and may be equal to 0.42 m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan-gearbox axialdistance may be in the range from 0.48 m to 0.8 m and optionally in therange from 0.50 m to 0.68 m (and may be equal to 0.58 m).

Fan Moment of Inertia

The fan 23 has a moment of inertial IF. The moment of inertia of the fanis measured based on the total mass of the rotor forming the fan, i.e.including the total mass of the plurality of fan blades, the fan hub andany support arm or other linkages provided to connect the fan to the fanshaft. The moment of inertia therefore includes all rotating componentsapart from the fan shaft. The moment of inertia is the mass moment ofinertia or rotational inertia of the fan with respect to rotation aroundthe principal rotational axis 9 of the engine. Rotation of the fan willcause a gyroscopic effect meaning that the fan shaft will tend tomaintain a steady direction of its axis of rotation. During maneuveringof the aircraft to which the gas turbine engine is mounted theorientation of the axis of rotation of the fan shaft will howeverchange. The gyroscopic effect will result in a reaction force at the fanshaft mounting structure to resist the tendency of the fan shaft tomaintain its orientation. The moment of inertia of the fan will have aneffect on the magnitude of the gyroscopic effect produced, and so has animpact on the design of the fan shaft mounting structure as is discussedelsewhere herein.

In various embodiments, the moment of inertia of the fan may be greaterthan or equal to 7.40×10⁷ kgm², and optionally greater than or equal to8.3×10⁷ kgm².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the moment of inertia ofthe fan may be greater than or equal to 7.4×10⁷ kgm² or 8.6×10⁷ kgm². Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the moment of inertia of the fan maybe greater than or equal to 3.0×10⁸ kgm² or 4.0×10⁸ kgm².

In various embodiments, the moment of inertia of the fan may be in therange from 7.40×10⁷ kgm² to 9.00×10⁸ kgm², and optionally in the rangefrom 8.3×10⁷ kgm² to 6.5×10⁸ kgm².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the moment of inertia ofthe fan may be in the range from 7.4×10⁷ kgm² to 1.5×10⁸ kgm² andoptionally in the range from 8.6×10⁷ kgm² to 9.6×10⁷ kgm² (and may beequal to 9.1×10⁷ kgm²).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the moment of inertia ofthe fan may be in the range from 3.0×10⁸ kgm² to 9.0×10⁸ kgm² andoptionally in the range from 4.0×10⁸ kgm² to 5.0×10⁸ kgm² (and may beequal to 4.5×10⁸ kgm²).

Power Transmitted by the Gearbox

Power is transmitted by the gearbox during operation of the engine. Thepower values given herein for the power transmitted by the gearbox arethe power transmitted by the gearbox at maximum take-off conditions. Themaximum take-off conditions are as defined elsewhere herein. The powertransmitted by the gearbox is defined as the power at the gearbox outputposition defined elsewhere herein.

In various embodiments, the power transmitted by the gearbox may begreater than or equal to 2.25×10⁷ W, and optionally greater than orequal to 2.5×10⁷ W.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, power transmitted by thegearbox may be greater than or equal to 2.25×10⁷ W or greater than orequal to 2.7×10⁷ W. In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the powertransmitted by the gearbox may be greater than or equal to 4.0×10⁷ W orgreater than or equal to 5.0×10⁷ W.

In various embodiments, the power transmitted by the gearbox may be inthe range from 2.25×10⁷ W to 1.00×10⁸ W, and optionally in the rangefrom 2.5×10⁷ W to 8.0×10⁷ W.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the power transmitted bythe gearbox may be in the range from 2.25×10⁷ W to 3.6×10⁷ W andoptionally in the range from 2.7×10⁷ W to 3.3×10⁷ W (and may be equal to3.0×10⁷ W).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the power transmitted bythe gearbox may be in the range from 4.0×10⁷ W to 1.0×10⁸ W andoptionally in the range from 5.0×10⁷ W to 6.0×10⁷ W (and may be equal to5.5×10⁷ W).

Carrier Torsional Stiffness

The torsional stiffness of the gearbox carrier is defined with referenceto FIGS. 27 to 31. FIG. 27 illustrates a close up view of the planetcarrier provided in the gearbox of any arrangement of the enginedescribed herein. As already described, the planet carrier 34 holds theplanet gears 32 in place. In planetary gearboxes 30, there may berelatively large centrifugal forces to be reacted by the carrier 34. Instar gearboxes, the centrifugal forces to react may be negligible orzero as there is no carrier rotation.

In the described arrangement, the planet carrier 34 comprises a forwardplate 34 a and a rearward plate 34 b, and pins 33 extending between theplates, as illustrated in FIG. 27. The pins 33 are arranged to beparallel to the engine axis 9. In alternative embodiments, a plate 34 bmay be provided on only one side—no plate or only a partial plate may beprovided on the other side of the carrier 34.

In the arrangement shown in FIG. 31, the carrier 34 additionallycomprises lugs 34 c, which may also be referred to as wedges or a web,between the forward and rearward plates 34 a, 34 b. The lugs 34 c mayincrease the overall stiffness of the carrier 34.

The stiffness of the carrier 34 is selected to be relatively high toreact centrifugal forces and/or to maintain gear alignment. The skilledperson would appreciate that stiffness is a measure of the displacementthat results from any applied forces or moments, and may not relate tostrength of the component. Hence to react a high load, any stiffness isacceptable so long as the resulting displacement is tolerable. How higha stiffness is desired to keep a displacement within acceptable limitstherefore depends on position and orientation of the gears, which isgenerally referred to as gear alignment (or mis-alignment).

Carrier torsional stiffness is a measure of the resistance of thecarrier 34 to an applied torque, τ, as illustrated in FIG. 28 (axialcross-section) and FIGS. 29 to 31 (a radial cross-section). The axis ofthe torque is parallel to the engine axis 9.

The diagonal lining of the plate 34 b at the rearward end of the carrier30 indicates that plate 34 b being treated as rigid and non-rotating (asfor a cantilever beam mounting). In embodiments with only one plate 34a, the ends of the pins 33 (and of the lugs 34 c if present) furtherfrom the single plate 34 a are held in place instead.

A torque, τ, is applied to the carrier 34 (at the position of the axialmid-point of the forward plate 34 a) and causes a rotationaldeformation, θ (e.g. twist) along the length of the carrier 34. Thetwist causes the carrier 34 to “wind up” as the ends of the pins 33 (andof the lugs 34 c if present) are held at a fixed radius on the carrierplates 34 a, 34 b.

The angle through which a point on an imaginary circle 902 on theforward plate 34 a passing through the rotation axis of each pin 33moves is θ where θ is the angle measured in radians. The imaginarycircle 902 may be referred to as the pin pitch circle diameter (pinPCD). The pin PCD may be in the range from 0.38 m to 0.65 m, for examplebeing equal to 0.4 m or 0.55 m. An effective linear torsional stiffnesscan therefore be defined for the carrier 34 as described above, usingthe radius r of the imaginary circle 902 (e.g. as illustrated in FIG.30).

In various embodiments, the torsional stiffness of the carrier 34 isgreater than or equal to 1.60×10⁸ Nm/rad, and optionally greater than orequal to 2.7×10⁸ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the torsional stiffness ofthe carrier 34 may be greater than or equal to 1.8×10⁸ Nm/rad, andoptionally may be greater than or equal to 2.5×10⁸ Nm/rad (andoptionally may be equal to 4.83×10⁸ Nm/rad). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the torsional stiffness of the carrier 34 may greaterthan or equal to 6.0×10⁸ Nm/rad and optionally may be greater than orequal to 1.1×10⁹ Nm/rad (and optionally may be equal to 2.17×10⁹Nm/rad).

In various embodiments, the torsional stiffness of the carrier 34 is inthe range from 1.60×10⁸ Nm/rad to 1.00×10¹¹ Nm/rad, and optionally inthe range from 2.7×10⁸ Nm/rad to 1×10¹⁰ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the torsional stiffness ofthe carrier 34 may be in the range from 1.8×10⁸ Nm/rad to 4.8×10⁹Nm/rad, and optionally may be in the range from 2.5×10⁸ Nm/rad to6.5×10⁸ Nm/rad (and optionally may be equal to 4.83×10⁸ Nm/rad). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the torsional stiffness of the carrier 34may be in the range from 6.0×10⁸ Nm/rad to 2.2×10¹⁰ Nm/rad andoptionally may be in the range from 1.1×10⁹ to 3.0×10⁹ Nm/rad (andoptionally may be equal to 2.17×10⁹ Nm/rad).

In various embodiments, the effective linear torsional stiffness of thecarrier 34 may be greater than or equal to 7.00×10⁹ N/m, and optionallygreater than or equal to 9.1×10⁹ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the effective lineartorsional stiffness of the carrier 34 may be greater than or equal to7.70×10⁹ N/m. In other such embodiments, the effective linear torsionalstiffness of the carrier 34 may be greater than or equal to 9.1×10⁹ N/m,optionally greater than or equal to 1.1×10¹⁰ N/m (and optionally may beequal to 1.26×10¹⁰ N/m). In some embodiments, for example in embodimentsin which the fan diameter is in the range from 330 to 380 cm, theeffective linear torsional stiffness of the carrier 34 may be greaterthan or equal to 1.2×10¹⁰ N/m and optionally may be greater than orequal to 2.1×10¹⁰ N/m (and optionally may be equal to 2.88×10¹⁰ N/m).

In various embodiments, the effective linear torsional stiffness of thecarrier 34 may be in the range from 7.00×10⁹ to 1.20×10¹¹ N/m, andoptionally in the range from 9.1×10⁹ N/m to 8.0×10¹⁰ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the effective lineartorsional stiffness of the carrier 34 may be in the range from 9.1×10⁹to 6.0×10¹⁰ N/m, and optionally may be in the range from 7×10⁹ N/m to2×10¹⁰ N/m, or from 8.5×10⁹ N/m to 2×10¹⁰ N/m (and optionally may beequal to 1.26×10¹⁰ N/m). In some embodiments, for example in embodimentsin which the fan diameter is in the range from 330 to 380 cm, theeffective linear torsional stiffness of the carrier 34 may be in therange from 1.2×10¹⁰ N/m to 1.2×10¹¹ N/m and optionally may be in therange from 1.0×10¹⁰ N/m to 5.0×10¹⁰ N/m (and optionally may be equal to2.88×10¹⁰ N/m).

The torsional stiffness of the carrier 34 may be controlled so as to bewithin a desired range by adjusting one or more parameters, includingcarrier material(s), carrier geometry, and the presence or absence oflugs.

Parameter Ratios

The inventor has discovered that the ratios (and/or products) of someproperties have a considerable impact on the operation of the gearboxand its inputs/outputs/support structure. Some or all of the below mayapply to any embodiment:

A system radial bending stiffness is defined by combining the radialbending stiffness of the fan shaft mounting structure 503 and the radialbending stiffness of the fan shaft 36 at the output of the gearbox inseries. The system radial bending stiffness is defined as:

$\frac{1}{( {1/K1} ) + ( {1/K2} )}$

Where K1 is the radial bending stiffness of the fan shaft mountingstructure, and K2 is radial bending stiffness of the fan shaft 36 at theoutput of the gearbox as defined elsewhere herein.

In various embodiments, the system radial bending stiffness may begreater than or equal to 3.90×10⁶ N/m and optionally greater than orequal to 3.6×10⁷ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the system radial bendingstiffness may be greater than or equal to 3.6×10⁷ N/m. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the system radial bending stiffness may begreater than or equal to 4.1×10⁷ N/m or greater than or equal to 5.8×10⁷N/m.

In various embodiments, the system radial bending stiffness may be inthe range from 3.90×10⁶ N/m to 1.40×10⁹ N/m, and optionally in the rangefrom 3.6×10⁷ N/m to 6.8×10⁸ N/m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the system radial bendingstiffness may be in the range from 3.6×10⁷ N/m to 4.0×10⁸ N/m andoptionally in the range from 3.6×10⁷ N/m to 4.9×10⁷ N/m (and may beequal to 3.9×10⁷ N/m).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the system radial bendingstiffness may be in the range from 4.1×10⁷ N/m to 1.4×10⁹ N/m, andoptionally in the range from 5.8×10⁷ N/m to 7.8×10⁷ N/m (and may beequal to 6.8×10⁷ N/m).

A fan shaft mounting radial bending stiffness ratio can be defined as:

$\frac{{the}{system}{radial}{bending}{stiffness}}{\begin{matrix}{{the}{radial}{bending}{stiffness}} \\{{of}{the}{fan}{shaft}{mounting}{structure}(503)}\end{matrix}}$

In various embodiments, the fan shaft mounting radial bending stiffnessratio may be greater than or equal to 1.0×10⁻³ optionally greater thanor equal to 5.0×10⁻³, or further optionally greater than or equal to2.0×10⁻².

In various embodiments, the fan shaft mounting radial bending stiffnessratio may be in the range from 1.0×10⁻³ to 4.0×10⁻¹, and optionally inthe range from 5.0×10⁻³ to 1.5×10⁻¹, in the range from 5.0×10⁻³ to2.0×10⁻² or in the range from 2.0×10⁻² to 1.5×10⁻¹.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft mountingradial bending stiffness ratio may be in the range from 2.2×10⁻² to3.2×10⁻² (and may be equal to 2.7×10⁻²).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft mountingradial bending stiffness ratio may be in the range from 2.6×10⁻² to3.6×10⁻² (and may be equal to 3.1×10⁻²).

In various embodiments, in addition to or alternatively to the fan shaftmounting radial bending stiffness ratio, a product of the parametersmaking up the fan shaft mounting radial bending stiffness ratio may bedefined. This product (referred to as the fan shaft mounting radialbending stiffness product) may be defined as:

the system radial bending stiffness×the radial bending stiffness of thefan shaft mounting structure (503)

In various embodiments, the fan shaft mounting radial bending stiffnessproduct may be greater than or equal to 2.7×10⁵ (N/m)², and optionallygreater than or equal to 4.0×10¹⁵ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft mountingradial bending stiffness product may be greater than or equal to4.3×10¹⁶ (N/m)². In some embodiments, for example in embodiments inwhich the fan diameter is in the range from 330 to 380 cm, the fan shaftmounting radial bending stiffness product may be greater than or equalto 4.3×10×10¹⁶ (N/m)².

In various embodiments, the fan shaft mounting radial bending stiffnessproduct may be in the range from 2.7×10¹⁵ (N/m)² to 9.0×10¹⁹ (N/m)², andoptionally in the range from 4.0×10¹⁵ (N/m)² to 1.5×10¹⁹ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft mountingradial bending stiffness product may be in the range from 4.3×10¹⁶ (N/m)to 3.0×10¹⁸ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft mountingradial bending stiffness product may be in the range from 4.3×10¹⁶(N/m)² to 9.0×10¹⁹ (N/m)².

A system tilt stiffness is defined by combining the tilt stiffness ofthe fan shaft mounting structure 503 and the tilt stiffness of the fanshaft 36 at the output of the gearbox in series. The system tiltstiffness is defined as:

$\frac{1}{( {1/K3} ) + ( {1/K4} )}$

Where K3 is the tilt stiffness of the fan shaft mounting structure, andK4 is the tilt stiffness of the fan shaft 36 at the output of thegearbox as defined elsewhere herein.

In various embodiments, the system tilt stiffness may be greater than orequal to 1.10×10⁵ Nm/rad and optionally greater than or equal to 8.5×10⁵Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the system tilt stiffnessmay be greater than or equal to 8.5×10⁵ Nm/rad. In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the system tilt stiffness may be greater than or equal to1.5×10⁶ Nm/rad or greater than or equal to 2.9×10⁶ Nm/rad.

In various embodiments, the system tilt stiffness may be in the rangefrom 1.10×10⁵ Nm/rad to 6.80×10⁷ Nm/rad, and optionally in the rangefrom 8.5×10⁵ Nm/rad to 3.4×10⁷ Nm/rad.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the system tilt stiffnessmay be in the range from 8.5×10⁵ Nm/rad to 1.2×10⁷ Nm/rad and optionallyin the range from 8.5×10⁵ Nm/rad to 1.7×10⁶ Nm/rad (and may be equal to1.2×10⁶ Nm/rad).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the system tilt stiffnessmay be in the range from 1.5×10⁶ Nm/rad to 6.8×10⁷ Nm/rad, andoptionally in the range from 2.9×10⁶ Nm/rad to 3.9×10⁶ Nm/rad (and maybe equal to 3.4×10⁶ Nm/rad).

In various embodiments, a fan shaft mounting tilt stiffness ratio isdefined as:

$\frac{{the}{system}{tilt}{stiffness}}{{tilt}{stiffness}{of}{the}{fan}{shaft}{mounting}{structure}(503)}$

In various embodiments, the fan shaft mounting tilt stiffness ratio maybe greater than or equal to 1.5×10⁻³, and optionally greater than orequal to 6.0×10⁻³ and further optionally greater than or equal to2.5×10⁻².

In various embodiments, the fan shaft mounting tilt stiffness ratio maybe in the range from 1.5×10⁻³ to 5.0×10⁻¹, and optionally in the rangefrom 6.0×10⁻³ to 2.0×10⁻¹, in the range from 6.0×10⁻³ to 2.5×10⁻² or inthe range from 2.5×10⁻² to 2.0×10⁻¹.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft mounting tiltstiffness ratio may be in the range from 3.0×10⁻² to 4.0×10⁻² (and maybe equal to 3.5×10⁻²).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft mounting tiltstiffness ratio may be in the range from 3.7×10⁻² to 4.7×10⁻² (and maybe equal to 4.2×10⁻²).

In various embodiments, in addition to or alternatively to the fan shaftmounting tilt stiffness ratio, a product of the parameters making upthat ratio may be defined. This product (referred to as the fan shaftmounting tilt stiffness product) may be defined as:

the system tilt stiffness×the tilt stiffness of the fan shaft mountingstructure (503)

In various embodiments, the fan shaft mounting tilt stiffness productmay be greater than or equal to 1.7×10¹² (Nm/rad)², and optionallygreater than or equal to 1.6×10¹³ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft mounting tiltstiffness product may be greater than or equal to 1.9×10¹³ (Nm/rad)². Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the fan shaft mounting tiltstiffness product may be greater than or equal to 3.0×10¹ (Nm/rad)².

In various embodiments, the fan shaft mounting tilt stiffness productmay be in the range from 1.7×10¹² (Nm/rad)² to 3.0×10¹⁷ (Nm/rad)², andoptionally in the range from 1.6×10¹³ (Nm/rad)² to 3.0×10¹⁶ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft mounting tiltstiffness product may be in the range from 1.9×10¹³ (Nm/rad)² to1.5×10¹⁶ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft mounting tiltstiffness product may be in the range from 3.0×10¹³ (Nm/rad)² to3.0×10¹⁷ (Nm/rad)².

A fan shaft radial bending stiffness ratio is defined as:

$\frac{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}(36){at}{the}{input}{to}{the}{fan}(23)}\end{matrix}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}{the}{fan}} \\{{shaft}(36){at}{the}{output}{of}{the}{gearbox}(30)}\end{matrix}}$

In various embodiments, the fan shaft radial bending stiffness ratio maybe greater than or equal to 6.0×10⁻³, and optionally greater than orequal to 0.015.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft radialbending stiffness ratio may be greater than or equal to 0.03 or greaterthan or equal to 0.07. In some embodiments, for example in embodimentsin which the fan diameter is in the range from 330 to 380 cm, the fanshaft radial bending stiffness ratio may be greater than or equal to0.02 or greater than or equal to 0.04.

In various embodiments, the fan shaft radial bending stiffness ratio maybe in the range from 6.0×10⁻³ to 2.5×10¹, and optionally in the rangefrom 0.015 to 2.5.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft radialbending stiffness ratio may be in the range from 0.03 to 0.85 andoptionally in the range from 0.07 to 0.27 (and may be equal to 0.17).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft radialbending stiffness ratio may be in the range from 0.02 to 1.5, andoptionally in the range from 0.04 to 0.24 (and may be equal to 0.14).

In various embodiments, in addition to or alternatively to the fan shaftradial bending stiffness ratio, a product of the parameters making upthat ratio may be defined. This product (referred to as the fan shaftradial bending stiffness product) may be defined as:

the radial bending stiffness of the fan shaft (36) at the input to thefan (23)×the radial bending stiffness of the fan shaft (36) at theoutput of the gearbox (30)

In various embodiments, the fan shaft radial bending stiffness productmay be greater than or equal to 1.2×10¹³ (N/m)², and optionally greaterthan or equal to 2.4×10¹⁴ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft radialbending stiffness product may be greater than or equal to 2.4×10¹⁴(N/m)². In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft radialbending stiffness product may be greater than or equal to 5.0×10¹³(N/m)².

In various embodiments, the fan shaft radial bending stiffness productmay be in the range from 1.2×10¹³ (N/m)² to 3.0×10¹⁸ (N/m)², andoptionally in the range from 2.4×10¹⁴ (N/m)² to 3.0×10¹⁷ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft radialbending stiffness product may be in the range from 2.4×10¹⁴ (N/m)² to2.7×10¹⁵ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft radialbending stiffness product may be in the range from 5.0×10¹³ (N/m)² to3.0×10¹⁸ (N/m)².

A fan shaft tilt stiffness ratio is defined as:

$\frac{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}(36)} \\{{at}{the}{input}{to}{the}{fan}(23)}\end{matrix}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}{fan}{shaft}(36)} \\{{at}{the}{output}{of}{the}{gearbox}(23)}\end{matrix}}$

In various embodiments, the fan shaft tilt stiffness ratio may begreater than or equal to 2.5×10⁻², and optionally greater than or equalto 0.05.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft tiltstiffness ratio may be greater than or equal to 0.2 or greater than orequal to 0.5. In some embodiments, for example in embodiments in whichthe fan diameter is in the range from 330 to 380 cm, the fan shaft tiltstiffness ratio may be greater than or equal to 0.1 or greater than orequal to 0.2.

In various embodiments, the fan shaft tilt stiffness ratio may be in therange from 2.5×10⁻² to 3.7×10², and optionally in the range from 0.05 to4.0×10¹.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft tiltstiffness ratio may be in the range from 0.2 to 5.0 and optionally inthe range from 0.5 to 1.5 (and may be equal to 1.00).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft tiltstiffness ratio may be in the range from 0.1 to 1.0×10¹, and optionallyin the range from 0.2 to 1.4 (and may be equal to 0.98).

In various embodiments, in addition to or alternatively to the fan shafttilt stiffness ratio, a product of the parameters making up that ratiomay be defined. This product (referred to as the fan shaft tiltstiffness product) may be defined as:

the tilt stiffness of the fan shaft (36) at the input to the fan(23)×the tilt stiffness of the fan shaft (36) at the output of thegearbox (30)

In various embodiments, the fan shaft tilt stiffness product may begreater than or equal to 3.5×10¹⁰ (Nm/rad)², and optionally greater thanor equal to 7.2×10¹¹ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft tiltstiffness product may be greater than or equal to 7.2×10¹¹ (Nm/rad)². Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the fan shaft tilt stiffness productmay be greater than or equal to 3.5×10¹¹ (Nm/rad)².

In various embodiments, the fan shaft tilt stiffness product may be inthe range from 3.5×10¹⁰ (Nm/rad)² to 5.0×10¹⁶ (Nm/rad)², and optionallyin the range from 7.2×10¹¹ (Nm/rad)² to 5.0×10¹⁵ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the fan shaft tiltstiffness product may be in the range from 7.2×10¹¹ (Nm/rad)² to1.5×10¹³ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the fan shaft tiltstiffness product may be in the range from 3.5×10¹¹ (Nm/rad)² to5.0×10¹⁶ (Nm/rad)².

In various embodiments, a first bearing separation ratio is defined as:

$\frac{{the}{first}{bearing}{separation}{distance}( d_{1} )}{\begin{matrix}{{the}{axial}{distance}{between}{the}{fan}{input}} \\{{position}{and}{the}{gearbox}{output}{{position}{}( d_{4} )}}\end{matrix}}$

In various embodiments, the first bearing separation ratio may begreater than or equal to 1.6×10⁻¹, and optionally greater than or equalto 1.8×10⁻¹, greater than or equal to 1.6×10⁻¹ or greater than or equalto 2.2×10⁻¹.

In various embodiments, the first bearing separation ratio may be in therange from 1.6×10⁻¹ to 3.3×10⁻¹, and optionally in the range from1.8×10⁻¹ to 3.0×10⁻¹, in the range from 1.6×10⁻¹ to 2.2×10⁻¹, or in therange from 2.2×10⁻¹ to 3.3×10⁻¹. The values in this and the previousparagraph may, for example, apply to embodiments in which the fandiameter is in the range from 240 to 280 cm or in the range from 330 cmto 380 cm.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm or in the range from 330 cmto 380 cm, the bearing separation ratio may be in the range from2.5×10⁻¹ to 2.9×10⁻¹ (and may be equal to 2.7×10⁻¹).

In various embodiments, in addition to or alternatively to the firstbearing separation ratio, a product of the parameters making up thatratio may be defined. This product (referred to as the first bearingseparation product) may be defined as:

the first bearing separation distance (d₁)×the axial distance betweenthe fan input position and the gearbox output position (d₄)

In various embodiments, the first bearing separation product may begreater than or equal to 5.2×10⁻² m², and optionally greater than orequal to 5.7×10⁻² m², or optionally greater than or equal to 7.5×10⁻²

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the first bearingseparation product may be greater than or equal to 5.2×10⁻² m². In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the first bearing separation product maybe greater than or equal to 7.5×10⁻² m².

In various embodiments, the first bearing separation product may be inthe range from 5.2×10⁻² m² to 2.6×10⁻¹ m², and optionally in the rangefrom 5.7×10⁻² m² to 2.4×10⁻¹ m², and optionally in the range from7.5×10⁻² m² to 2.6×10⁻¹ m²

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the first bearingseparation product may be in the range from 5.2×10⁻² m² to 1.4×10⁻¹ m².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the first bearingseparation product may be in the range from 7.5×10⁻² m² to 2.6×10⁻¹ m².

For embodiments in which there is at least a first and second supportingbearing forward of the gearbox (and rearward of the fan input position)a second bearing separation ratio is defined as:

$\frac{{the}{first}{bearing}{seperation}{{distance}{}( d_{1} )}}{{the}{bearing}{axial}{seperation}( d_{3} )}$

In various embodiments, the second bearing separation ratio may begreater than or equal to 4.1×10⁻¹, and optionally greater than or equalto 4.5×10⁻¹, and further optionally greater than or equal to 6.0×10⁻¹.

In various embodiments, the second bearing separation ratio may be inthe range from 4.1×10⁻¹ to 8.3×10⁻¹, and optionally in the range from4.5×10⁻¹ to 7.7×10⁻¹, or in the range from 4.1×10⁻¹ to 6.0×10⁻¹, or inthe range 6.0×10⁻¹ to 8.3×10⁻¹. The values in this and the previousparagraph may, for example, apply to embodiments in which the fandiameter is in the range from 240 to 280 cm or in the range from 330 cmto 380 cm.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm or in the range from 330 cmto 380 cm, or in other various embodiments, the second bearingseparation ratio may be in the range from 6.4×10⁻¹ to 7.0×10⁻¹ (and maybe equal to 6.7×10⁻¹).

In various embodiments, in addition to or alternatively to the secondbearing separation ratio, a product of the parameters making up thatratio may be defined. This product (referred to as the second bearingseparation product) may be defined as:

the first bearing separation distance (d₁)×the bearing axial separation(d₃)

In various embodiments, the second bearing separation product may begreater than or equal to 2.0×10⁻² m², and optionally greater than orequal to 2.3×10⁻² m² and further optionally greater than or equal to3.5×10⁻² m².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the second bearingseparation product may be greater than or equal to 2.0×10⁻² m². In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the second bearing separation product maybe greater than or equal to 3.5×10⁻² m.

In various embodiments, the second bearing separation product may be inthe range from 2.0×10⁻² m² to 1.1×10⁻¹ m², and optionally in the rangefrom 2.3×10⁻² m² to 8.5×10⁻² m², or may be in the range from 3.5×10⁻² m²to 1.1×10⁻¹ m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the second bearingseparation product may be in the range from 2.0×10⁻² m² to 5.6×10⁻² m.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the second bearingseparation product may be in the range from 3.5×10⁻² m² to 1.1×10⁻¹ m.

In various embodiments, a first planet carrier stiffness ratio may bedefined as:

$\frac{\begin{matrix}{{the}{effective}{linear}{torsional}} \\{{stiffness}{of}{the}{planet}{carrier}(34)}\end{matrix}}{\begin{matrix}{{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}(503)}\end{matrix}}$

In various embodiments, the first planet carrier stiffness ratio may begreater than or equal to 7.0×10⁻³, and optionally greater than or equalto 7.0×10⁻².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the first planet carrierstiffness ratio may be greater than or equal to 6.9 or may be greaterthan or equal to 7.0 or 4.0. In some embodiments, for example inembodiments in which the fan diameter is in the range from 330 to 380cm, the first planet carrier stiffness ratio may be greater than orequal to 7.0 or greater than or equal to 1.0×10¹ or 5.0

In various embodiments, the first planet carrier stiffness ratio may bein the range from 7.0×10⁻³ to 1.9×10³, and optionally in the range from7.0×10⁻² to 9.0×10¹.

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the first planet carrierstiffness ratio may be in the range from 6.9 to 1.2×10² and optionallyin the range from 7.0 to 11.0 (and may be equal to 8.5). In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the first planet carrier stiffness ratiomay be in the range from in the range from 4.0 to 6.0 (and may be equalto 5.2)

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the first planet carrierstiffness ratio may be in the range from 7.0 to 1.9×10³ and optionallyin the range from 1.0×10¹ to 2.0×10¹ (and may be equal to 1.33×10¹). Insome embodiments, for example in embodiments in which the fan diameteris in the range from 330 to 380 cm, the first planet carrier stiffnessratio may be in the range from 5.0 to 9.0 (and may be equal to 7.1).

In various embodiments, a first planet carrier stiffness product may bedefined as:

(the effective linear torsional stiffness of the planet carrier(34))×(radial bending stiffness of the fan shaft mounting structure(503))

In various embodiments, the first planet carrier stiffness product maybe greater than or equal to 2.9×10¹⁸ (N/m)², and optionally greater thanor equal to 5.0×10¹⁸ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the first planet carrierstiffness product may be greater than or equal to 8.0×10¹⁸ (N/m)² orgreater than or equal to 9.0×10¹⁸ (N/m)². In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the first planet carrier stiffness product may be greaterthan or equal to 1.0×10¹⁹ (N/m)² or 2.0×10¹⁹ (N/m)² or 5.0×10¹⁹ (N/m)².

In various embodiments, the first planet carrier stiffness product maybe in the range from 2.9×10¹⁸ (N/m)² to 8.0×10²² (N/m)², and optionallyin the range from 5.0×10¹⁸ (N/m)² to 8.0×10²¹ (N/m)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the first planet carrierstiffness product may be in the range from 8.0×10¹⁸ (N/m)² to 8.0×10²¹(N/m)² and optionally in the range from 9.0×10¹⁸ (N/m)² to 3.0×10¹⁹(N/m)² (and may be equal to 1.1×10¹⁹ (N/m)² or may be equal to 1.9×10¹⁹(N/m)²).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the first planet carrierstiffness product may be in the range from 1.0×10¹⁹ (N/m)² to 8.0×10²²(N/m)² and optionally in the range from 2.0×10¹⁹ (N/m)² to 4.0×10¹⁹(N/m)² (and may be equal to 3.3×10¹⁹ (N/m)²), or optionally in the rangefrom 5.0×10¹⁹ (N/m)² to 8.0×10¹⁹ (N/m²) (and may be equal to 6.2×10¹⁹(N/m)²).

In various embodiments, a second planet carrier stiffness ratio may bedefined as:

$\frac{\begin{matrix}{{the}{torsional}} \\{{stiffness}{of}{the}{planet}{carrier}(34)}\end{matrix}}{\begin{matrix}{{the}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}(503)}\end{matrix}}$

In various embodiments, the second planet carrier stiffness ratio may begreater than or equal to 6.0×10⁻³, and optionally greater than or equalto 6.0×10⁻².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the second planet carrierstiffness ratio may be greater than or equal to 1.36×10¹ or 7.9. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 330 to 380 cm, the second planet carrier stiffness ratiomay be greater than or equal to 1.5×10¹ or greater than or equal to1.7×10¹ or 0.4×10¹.

In various embodiments, the second planet carrier stiffness ratio may bein the range from 6.0×10⁻³ to 7.0×10³, and optionally in the range from6.0×10⁻² to 7.0×10².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the second planet carrierstiffness ratio may be in the range from 1.36×10¹ to 7.0×10² andoptionally in the range from 1.36×10¹ to 2.4×10¹ (and may be equal to1.4×10¹). In some embodiments, for example in embodiments in which thefan diameter is in the range from 240 to 280 cm, the second planetcarrier stiffness ratio may be in the range from 7.9 to 9.9 (and may beequal to 8.9)

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the second planet carrierstiffness ratio may be in the range from 1.5×10¹ to 7.0×10³ andoptionally in the range from 1.7×10¹ to 3.7×10¹ (and may be equal to2.7×10¹).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the second planet carrierstiffness ratio may in the range from 0.4×10¹ to 2.4×10¹ (and may beequal to 1.4×10¹).

In various embodiments, a second planet carrier stiffness product may bedefined as:

(the torsional stiffness of the planet carrier (34))×(tilt stiffness ofthe fan shaft mounting structure (503))

In various embodiments, the second planet carrier stiffness product maybe greater than or equal to 2.4×10⁵ (Nm/rad)², and optionally greaterthan or equal to 4.9×10⁵ (Nm/rad)² In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the second planet carrier stiffness product may be greater than orequal to 4.9×10¹⁵ (Nm/rad)² or 7.9×10¹⁵ (Nm/rad)² or 1.0×10¹⁶ (Nm/rad)².In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the second planet carrierstiffness product may be greater than or equal to 9.0×10¹⁵ (Nm/rad)² or7.4×10¹⁶ (Nm/rad)² or 1.0×10¹⁷ (Nm/rad)².

In various embodiments, the second planet carrier stiffness product maybe in the range from 2.4×10¹⁵ (Nm/rad)² to 2.7×10²¹ (Nm/rad)², andoptionally in the range from 4.9×10¹⁵ (Nm/rad)² to 2.0×10²⁰ (Nm/rad)².

In some embodiments, for example in embodiments in which the fandiameter is in the range from 240 to 280 cm, the second planet carrierstiffness product may be in the range from 4.9×10¹⁵ (Nm/rad)² to9.0×10¹⁹ (Nm/rad)² and optionally in the range from 7.9×10¹⁵ (Nm/rad)²to 1.2×10¹⁶ (Nm/rad)² (and may be equal to 9.9×10¹⁵ (Nm/rad)²), oroptionally in the range from 1.0×10¹⁶ (Nm/rad)² to 1.2×10¹⁶ (Nm/rad)²(and may be equal to 1.6×10¹⁶ (Nm/rad)²).

In some embodiments, for example in embodiments in which the fandiameter is in the range from 330 to 380 cm, the second planet carrierstiffness product may be in the range from 9.0×10¹⁵ (Nm/rad)² to2.7×10²¹ (Nm/rad)² and optionally in the range from 7.4×10¹⁶ (Nm/rad)²to 1.1×10¹⁷ (Nm/rad)² (and may be equal to 9.4×10¹⁶ (Nm/rad)²), oroptionally in the range from 1.0×10¹⁷ (Nm/rad)² to 2.6×10¹⁷ (Nm/rad)²(and may be equal to 1.8×10¹⁷ (Nm/rad)²).

FIG. 32 illustrates an example aircraft 1000 having a gas turbine engine10 attached to each wing 1002 a, 1002 b thereof. Each gas turbine engine10 is attached via a respective pylon 1004 a, 1004 b. The gas turbines10 may be that of any embodiment described herein. The aircraft shown inFIG. 32 is to be understood as the aircraft for which the gas turbineengine 10 of any embodiment or aspect disclosed herein has been designedto be attached. The aircraft 1000 has a cruise condition correspondingto the cruise conditions defined elsewhere herein and a max take-offcondition corresponding to the MTO conditions defined elsewhere herein.

The present disclosure also relates to a method 2000 of operating a gasturbine engine on an aircraft (e.g. the aircraft of FIG. 32). The method2000 is illustrated in FIG. 33. The method 2000 comprises operating 2010the gas turbine engine 10 described elsewhere herein to providepropulsion for the aircraft to which it is mounted under maximumtake-off conditions. The method further comprises operating 2020 the gasturbine engine to provide propulsion during cruise conditions. The gasturbine engine is operated such that any of the parameters or ratiosdefined herein are within the specified ranges. Cruise conditions andmax-take off conditions are as defined elsewhere herein.

The torque on the core shaft 26 may be referred to as the input torque,as this is the torque which is input to the gearbox 30. The torquesupplied by the turbine 19 to the core shaft (i.e. the torque on thecore shaft) at cruise conditions may be greater than or equal to 10,000Nm, and optionally greater than or equal to 11,000 Nm. In someembodiments, for example in embodiments in which the fan diameter is inthe range from 240 to 280 cm, the torque on the core shaft 26 at cruiseconditions may be greater than or equal to 10,000 or 11,000 Nm (andoptionally may be equal to 12,760 Nm). In some embodiments, for examplein embodiments in which the fan diameter is in the range from 330 to 380cm, the torque on the core shaft 26 at cruise conditions may be greaterthan or equal to 25,000 Nm, and optionally greater than or equal to30,000 Nm (and optionally may be equal to 34,000 Nm).

The torque on the core shaft at cruise conditions may be in the rangefrom 10,000 to 50,000 Nm, and optionally from 11,000 to 45,000 Nm. Insome embodiments, for example in embodiments in which the fan diameteris in the range from 240 to 280 cm, the torque on the core shaft 26 atcruise conditions may be in the range from 10,000 to 15,000 Nm, andoptionally from 11,000 to 14,000 Nm (and optionally may be equal to12,760 Nm). In some embodiments, for example in embodiments in which thefan diameter is in the range from 330 to 380 cm, the torque on the coreshaft 26 at cruise conditions may be in the range from 25,000 Nm to50,000 Nm, and optionally from 30,000 to 40,000 Nm (and optionally maybe equal to 34,000 Nm).

Under maximum take-off (MTO) conditions, the torque on the core shaft 26may be greater than or equal to 28,000 Nm, and optionally greater thanor equal to 30,000 Nm. In some embodiments, for example in embodimentsin which the fan diameter is in the range from 240 to 280 cm, the torqueon the core shaft 26 under MTO conditions may be greater than or equalto 28,000, and optionally greater than or equal to 35,000 Nm (andoptionally may be equal to 36,300 Nm). In some embodiments, for examplein embodiments in which the fan diameter is in the range from 330 to 380cm, the torque on the core shaft 26 under MTO conditions may greaterthan or equal to 70,000 Nm, and optionally greater than or equal to80,000 or 82,000 Nm (and optionally may be equal to 87,000 Nm).

Under maximum take-off (MTO) conditions, the torque on the core shaft 26may be in the range from 28,000 Nm to 135,000 Nm, and optionally in therange from 30,000 to 110,000 Nm. In some embodiments, for example inembodiments in which the fan diameter is in the range from 240 to 280cm, the torque on the core shaft 26 under MTO conditions may be in therange from 28,000 to 50,000 Nm, and optionally from 35,000 to 38,000 Nm(and optionally may be equal to 36,300 Nm). In some embodiments, forexample in embodiments in which the fan diameter is in the range from330 to 380 cm, the torque on the core shaft 26 under MTO conditions maybe in the range from 70,000 Nm to 135,000 Nm, and optionally from 80,000to 90,000 Nm or 82,000 to 92,000 Nm (and optionally may be equal to87,000 Nm).

Torque has units of [force]×[distance] and may be expressed in units ofNewton metres (N·m), and is defined in the usual way as would beunderstood by the skilled person.

FIG. 34 illustrates how the stiffnesses defined herein may be measured.FIG. 34 shows a plot of the displacement 6 resulting from theapplication of a load L (e.g. a force, moment or torque) applied to acomponent for which the stiffness is being measured. At levels of loadfrom zero to L_(G) there is a non-linear region in which displacement iscaused by motion of the component (or relative motion of separate partsof the component) as it is loaded, rather than deformation of thecomponent; for example moving within clearance between parts. At levelsof load above L_(H) the elastic limit of the component has been exceededand the applied load no longer causes elastic deformation—plasticdeformation or failure of the component may occur instead. Betweenpoints G and H the applied load and resulting displacement have a linearrelationship. The stiffnesses defined herein may be determined bymeasuring the gradient of the linear region between points G and H (withthe stiffness being the inverse of that gradient). The gradient may befound for as large a region of the linear region as possible to increasethe accuracy of the measurement by providing a larger displacement tomeasure. For example, the gradient may be found by applying a load equalto or just greater than L_(G) and equal to or just less than L_(H).Although the displacement is referred to as δ in this description, theskilled person would appreciate that equivalent principles would applyto a linear or angular displacement.

The stiffnesses defined herein, unless otherwise stated, are for thecorresponding component(s) when the engine is off (i.e. at zero speed/onthe bench) The stiffnesses generally do not vary over the operatingrange of the engine; the stiffness at cruise conditions of the aircraftto which the engine is used (those cruise conditions being as definedelsewhere herein) may therefore be the same as for when the engine isnot in use. However, where the stiffness varies over the operating rangeof the engine, the stiffnesses defined herein are to be understood asbeing values for when the engine is at room temperature and unmoving.

Anything described herein with reference to a planetary type gearbox canapply equally to a star type gearbox unless otherwise stated or where itis apparent that a feature is specific to a particular gearbox type.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine for an aircraft comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades, the fan having a fan axialcentreline; a gearbox that is configured to: receive an input from thecore shaft, and output drive to a fan shaft via an output of the gearboxso as to drive the fan at a lower rotational speed than the core shaft;and a fan shaft mounting structure arranged to mount the fan shaftwithin the engine, the fan shaft mounting structure comprising at leasttwo supporting bearings connected to the fan shaft, wherein: the fancomprises 22, 24 or 26 fan blades; a system radial bending stiffness isdefined as: $\frac{1}{\begin{matrix}{( \frac{1}{\begin{matrix}{a{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) +} \\( \frac{1}{\begin{matrix}{a{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}} )\end{matrix}};$ a system tilt stiffness is defined as:$\frac{1}{\begin{matrix}{( \frac{1}{\begin{matrix}{a{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) +} \\( \frac{1}{\begin{matrix}{a{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}} )\end{matrix}};$ and either: (i) the system radial bending stiffness isgreater than or equal to 3.90×10⁶ N/m; or (ii) the system tilt stiffnessis greater than or equal to 1.10×10⁵ Nm/rad.
 2. The gas turbine engineaccording to claim 1, wherein a fan shaft mounting radial bendingstiffness ratio of:$\frac{{the}{system}{radial}{bending}{stiffness}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}{the}} \\{{fan}{shaft}{mounting}{structure}}\end{matrix}}$ is in the range from 1.0×10⁻³ to 4.0×10⁻¹.
 3. The gasturbine engine according to claim 1, wherein a fan shaft mounting tiltstiffness ratio of:$\frac{{the}{system}{tilt}{stiffness}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}} \\{{fan}{shaft}{mourting}{structure}}\end{matrix}}$ is in the range from 1.5×10⁻³ to 5.0×10⁻¹.
 4. The gasturbine engine according to claim 1, wherein at least one of thefollowing is satisfied: a) the system radial bending stiffness is in arange from 3.90×10⁶ N/m to 1.40×10⁹ N/m; b) the radial bending stiffnessof the fan shaft mounting structure is greater than or equal to 7.00×10⁸N/m; and c) the radial bending stiffness of the fan shaft at the outputof the gearbox is greater than or equal to 4.00×10⁶ N/m.
 5. The gasturbine engine according to claim 1, wherein at least one of thefollowing is satisfied: a) the gearbox has a gear ratio of between 3 and3.1; b) an axial distance between an input to the fan and a closestbearing of the at least two supporting bearings in a rearward directionfrom the fan is in a range from 0.12 m to 0.40 m; c) a fan tip loadingis configured to be in a range from 0.30 to 0.34 at cruise conditions,where the fan tip loading is defined as dH/U_(tip) ², where dH is anenthalpy rise across the fan and U_(tip) is a translational velocity ofa fan tip of the fan; and d) the at least two supporting bearingscomprise a first supporting bearing and second supporting bearing,wherein both of the first and the second supporting bearings are locatedat positions forward of the gearbox.
 6. The gas turbine engine accordingto claim 1, wherein at least one of the following is satisfied: a) thefan has a diameter in a range of 220 cm to 240 cm; b) a first bearingseparation product defined as:a first bearing separation distance×an axial distance between a faninput position and a gearbox output position is in a range from 5.2×10⁻²m² to 2.6×10⁻¹ m²; c) the fan blades are manufactured using a titaniumbased metal; and d) a temperature of a flow at an exit to the combustoris configured to be greater than 1600K at cruise conditions.
 7. A gasturbine engine for an aircraft comprising: an engine core comprising aturbine, a compressor, and a core shaft connecting the turbine to thecompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades, the fan having a fan axialcentreline; a gearbox that is configured to: receive an input from thecore shaft, and output drive to a fan shaft via an output of the gearboxso as to drive the fan at a lower rotational speed than the core shaft;and a fan shaft mounting structure arranged to mount the fan shaftwithin the engine, the fan shaft mounting structure comprising at leasttwo supporting bearings connected to the fan shaft, wherein: the fanblades are manufactured using a titanium based metal; a system radialbending stiffness is defined as: $\frac{1}{\begin{matrix}{( \frac{1}{\begin{matrix}{a{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) +} \\( \frac{1}{\begin{matrix}{a{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}} )\end{matrix}};$ a system tilt stiffness is defined as:$\frac{1}{\begin{matrix}{( \frac{1}{\begin{matrix}{a{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) +} \\( \frac{1}{\begin{matrix}{a{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}} )\end{matrix}};$ and either: (i) the system radial bending stiffness isgreater than or equal to 3.90×10⁶ N/m; or (ii) the system tilt stiffnessis greater than or equal to 1.10×10⁵ Nm/rad.
 8. The gas turbine engineaccording to claim 7, wherein a fan shaft mounting radial bendingstiffness ratio of:$\frac{{the}{system}{radial}{bending}{stiffness}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}{the}} \\{{fan}{shaft}{mounting}{structure}}\end{matrix}}$ is in a range from 1.0×10⁻³ to 4.0×10⁻¹.
 9. The gasturbine engine according to claim 7, wherein a fan shaft mounting tiltstiffness ratio of:$\frac{{the}{system}{tilt}{stiffness}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}} \\{{fan}{shaft}{mounting}{structure}}\end{matrix}}$ is in a range from 1.5×10⁻³ to 5.0×10⁻¹.
 10. The gasturbine engine according to claim 7, wherein at least one of thefollowing is satisfied: a) the system radial bending stiffness is in arange from 3.90×10⁶ N/m to 1.40×10⁹ N/m; b) the radial bending stiffnessof the fan shaft mounting structure is greater than or equal to 7.00×10⁸N/m; and c) the radial bending stiffness of the fan shaft at the outputof the gearbox is greater than or equal to 4.00×10⁶ N/m.
 11. The gasturbine engine according to claim 7, wherein at least one of thefollowing is satisfied: a) the gearbox has a gear ratio of between 3 and3.1; b) an axial distance between an input to the fan and a closestbearing of the at least two supporting bearings in a rearward directionfrom the fan is in a range from 0.12 m to 0.40 m; c) a fan tip loadingis configured to be in a range from 0.30 to 0.34 at cruise conditions,where the fan tip loading is defined as dH/U_(tip) ², where dH is anenthalpy rise across the fan and U_(tip) is a translational velocity ofa fan tip of the fan; and d) the at least two supporting bearingscomprise a first supporting bearing and second supporting bearing,wherein both of the first and the second supporting bearings are locatedat positions forward of the gearbox.
 12. The gas turbine engineaccording to claim 7, wherein at least one of the following issatisfied: a) the fan has a diameter in a range of 220 cm to 240 cm; andb) a first bearing separation product defined as:a first bearing separation distance×an axial distance between a faninput position and a gearbox output position is in a range from 5.2×10⁻²m² to 2.6×10⁻¹ m²; c) the fan comprises 22, 24 or 26 fan blades; and d)a temperature of a flow at an exit to the combustor is configured to begreater than 1600K at cruise conditions.
 13. A gas turbine engine for anaircraft comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades, the fan having a fan axial centreline; a gearbox that isconfigured to: receive an input from the core shaft, and output drive toa fan shaft via an output of the gearbox so as to drive the fan at alower rotational speed than the core shaft; and a fan shaft mountingstructure arranged to mount the fan shaft within the engine, the fanshaft mounting structure comprising at least two supporting bearingsconnected to the fan shaft, wherein: a temperature of a flow at an exitto the combustor is configured to be greater than 1600K at cruiseconditions; a system radial bending stiffness is defined as:$\frac{1}{\begin{matrix}{( \frac{1}{\begin{matrix}{a{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) +} \\( \frac{1}{\begin{matrix}{a{radial}{bending}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}} )\end{matrix}};$ a system tilt stiffness is defined as:$\frac{1}{\begin{matrix}{( \frac{1}{\begin{matrix}{a{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{mounting}{structure}}\end{matrix}} ) +} \\( \frac{1}{\begin{matrix}{a{tilt}{stiffness}{of}} \\{{the}{fan}{shaft}{at}{the}{output}{of}{the}{gearbox}}\end{matrix}} )\end{matrix}};$ and either: (i) the system radial bending stiffness isgreater than or equal to 3.90×10⁶ N/m; or (ii) the system tilt stiffnessis greater than or equal to 1.10×10⁵ Nm/rad.
 14. The gas turbine engineaccording to claim 13, wherein a fan shaft mounting radial bendingstiffness ratio of:$\frac{{the}{system}{radial}{bending}{stiffness}}{\begin{matrix}{{the}{radial}{bending}{stiffness}{of}{the}} \\{{fan}{shaft}{mounting}{structure}}\end{matrix}}$ is in a range from 1.0×10⁻³ to 4.0×10⁻¹.
 15. The gasturbine engine according to claim 13, wherein a fan shaft mounting tiltstiffness ratio of:$\frac{{the}{system}{tilt}{stiffness}}{\begin{matrix}{{the}{tilt}{stiffness}{of}{the}} \\{{fan}{shaft}{mounting}{structure}}\end{matrix}}$ is in a range from 1.5×10⁻³ to 5.0×10⁻¹.
 16. The gasturbine engine according to claim 13, wherein at least one of thefollowing is satisfied: a) the system radial bending stiffness is in arange from 3.90×10⁶ N/m to 1.40×10⁹ N/m; b) the radial bending stiffnessof the fan shaft mounting structure is greater than or equal to 7.00×10⁸N/m; and c) the radial bending stiffness of the fan shaft at the outputof the gearbox is greater than or equal to 4.00×10⁶ N/m.
 17. The gasturbine engine according to claim 13, wherein at least one of thefollowing is satisfied: a) the gearbox has a gear ratio of between 3 and3.1; b) an axial distance between an input to the fan and a closestbearing of the at least two supporting bearings in a rearward directionfrom the fan is in a range from 0.12 m to 0.40 m; c) a fan tip loadingis configured to be in a range from 0.30 to 0.34 at cruise conditions,where the fan tip loading defined as dH/U_(tip) ², where dH is anenthalpy rise across the fan and U_(tip) is a translational velocity ofa fan tip of the fan; d) the at least two supporting bearings comprise afirst supporting bearing and second supporting bearing, wherein both ofthe first and the second supporting bearings are located at positionsforward of the gearbox.
 18. The gas turbine engine according to claim13, wherein at least one of the following is satisfied: a) the fan has adiameter in a range of 220 cm to 240 cm; b) a first bearing separationproduct defined as:a first bearing separation distance×an axial distance between a faninput position and a gearbox output position is in a range from 5.2×10⁻²m² to 2.6×10⁻¹ m²; c) the fan comprises 22, 24 or 26 fan blades; and d)the fan blades are manufactured using a titanium based metal.